Advanced design for lightweight structures: Review and prospects

Progress in Aerospace Sciences ∎ (∎∎∎∎) ∎∎∎–∎∎∎

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Advanced design for lightweight structures: Review and prospects Daniel F.O. Braga a,n, S.M.O. Tavares a, Lucas F.M. da Silva b,c, P.M.G.P. Moreira a, Paulo M.S. T. de Castro c a

Institute of Mechanical Engineering and Industrial Management (INEGI), Universidade do Porto, Portugal Institute of Mechanical Engineering (IDMEC), Universidade do Porto, Portugal c Faculdade de Engenharia da Universidade do Porto (FEUP), Portugal b

art ic l e i nf o

a b s t r a c t

Article history: Received 13 December 2013 Received in revised form 12 March 2014 Accepted 25 March 2014

Current demand for fuel efficient aircraft has been pushing the aeronautical sector to develop ever more lightweight designs while keeping safe operation and required structural strength. Along with lightweighting, new structural design concepts have also been established in order to maintain the aircraft in service for longer periods of time, with high reliability levels. All these innovations and requirements have led to deeply optimized aeronautical structures contributing to more sustainable air transport. This article reviews the major design philosophies which have been employed in aircraft structures, including safe-life, fail-safe and damage tolerance taking into account their impact on the structural design. A brief historical review is performed in order to analyse what led to the development of each philosophy. Material properties are related to each of the design philosophies. Damage tolerant design has emerged as the main structural design philosophy in aeronautics, requiring deep knowledge on materials fatigue and corrosion strength, as well as potential failure modes and non-destructive inspection techniques, particularly minimum detectable defect and scan times. A discussion on the implementation of structural health monitoring and self-healing structures within the current panorama of structures designed according to the damage tolerant philosophy is presented. This discussion is aided by a review of research on these two subjects. These two concepts show potential for further improving safety and durability of aircraft structures. & 2014 Elsevier Ltd. All rights reserved.

Keywords: Structural design Safe-life Fail-safe Damage tolerance Structural health monitoring Self-healing structures

Contents 1. 2. 3.

Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Historical background . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Structural design philosophy. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 3.1. Safe-life design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 3.2. Fail-safe concepts. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4 3.3. Damage tolerant design. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4 4. Structural health monitoring. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6 5. Self-healing structures. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8 6. Concluding remarks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9 Acknowledgements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10 References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

1. Introduction

n

Corresponding author. E-mail address: [email protected] (D.F.O. Braga).

Many definitions for structural design may be found in the literature. One holistic definition is given by the McGraw-Hill Concise Encyclopaedia of Engineering [1] stating that structural design is the science of “selection of materials and member type,

http://dx.doi.org/10.1016/j.paerosci.2014.03.003 0376-0421/& 2014 Elsevier Ltd. All rights reserved.

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Acronyms ASIP DCB FBG FSW IVHM LBW

aircraft structural integrity programme double cantilever beam fibre Bragg grating friction stir welding integrated vehicle health management laser beam welding

size, and configuration to carry loads in a safe and serviceable fashion…”. Further, [1] states that structural design normally includes at least five distinct phases – project requirements, materials, structural scheme, analysis, and design, although in many structures one more stage may be required, which is testing. More complex structures, which make use of more state-of-art materials or concepts, regularly require proof of concept and demonstration of their capabilities and longevity. In summary, structural design may be described as the combination of several realms of engineering to the process of new product development. These realms include, mechanical, civil, and materials engineering among others. Over the past few decades, structural design for aeronautical structures has been subjected to multiple pressures of different stakeholders in order to decrease their footprint and improve their safety for a more affordable and more competitive transportation of goods and people. This is summarized by “Quieter, Cleaner and Greener” or “More Affordable, Cleaner and Quieter” drivers which guide the development of the civil aeronautical industry [2]. One example of this effort is the reduction of emissions in this industry by the European Union Clean Sky initiative [3], which aims at reducing 50% CO2 emissions and 80% NOx emissions by 2020. Civil aircraft structures are one of the most efficient compared with other vehicles structures whilst at the same time presenting high safety records due to the constant innovation and the usage of advanced design principles. The design principles employed to develop aeronautical structures explore all properties and failure modes of the materials in order to create optimal structures which tolerate inherent imperfections. Although these design concepts are regularly discussed, several experts present slightly different interpretations. The importance of using correct language for aeronautical fatigue is addressed in [4]. All the stakeholders that operate in this field are responsible for the eventual misunderstandings generated, from the different regulations issued by the airworthiness authorities and their application by aircraft manufacturers and airline operators, to the interpretation of maintenance plans by the aircraft operators. This article focuses on the three main design philosophies applied to high performance structures, safe-life, fail-safe and damage tolerant design, highlighting the differences between them and their respective scope of application. Each philosophy is discussed in terms of material properties, highlighting the key material properties for each design methodology. Limitations of safe-life design methodology, leading to its replacement by damage tolerant design, are discussed. A final topic regarding the effect of newer concepts on design philosophy, such as structural health monitoring and selfhealing structures are addressed. A review on these two concepts is presented and used to show their potential for improvement in maintaining continuous airworthiness of structures.

2. Historical background In the early days of flight, aircrafts were designed solely on the basis of static strength. At the time this was sufficient, as airplanes

MSD NDI POD SHM USAF WFD WSPS X-FEM

multiple site damage non-destructive inspection probability of detection structural health monitoring United States air force widespread fatigue damage weapon system performance specification extended finite elements method

were not able to perform long distance flights and had a short lifetime. Factors such as fatigue, corrosion, accidental damage, among others, were not taken into account. As initial aviation evolved from a mere hobby or technological demonstration to a serious mean of transportation or warfare, especially with requirements imposed by the two world wars, structural integrity became a relevant issue. The first implemented design approach was the safe-life or safety-by-retirement, where a structure is operated during a service life with a low probability of failure, being retired at the end of this safe life (which is the predicted safe-life plus an extra safety margin to take into account the uncertainty). As during war time aircrafts become obsolete before the safe life is over, this design methodology allowed for very safe structures. Materials fatigue became a more prominent issue with the start of the jet age, as commercial airplanes aimed at longer distances and at higher altitudes, increasing the applied loads (e.g. cabin pressurization). The Havilland Comet crashes [5] have shown the limitations of this philosophy, as fatigue cracks occurred earlier than anticipated. The safe-life of the aircraft was determined through an experimental programme, but unaccounted phenomena have led to a non-conservative estimate. The Comet crashes revealed limitations in the fatigue analyses, leading to the conclusion that safety could not be guaranteed on a safe-life basis without imposing uneconomically short service lives on major components of the structure. Fail-safe was put forward to address these limitations. This new concept involved designing a structure which could sustain a satisfactory life span without damage, but also allowed for inspectability and multiple load paths, in order to avoid complete structural failure within service. The last point is associated with the concept of residual strength which was introduced along with fail-safe design. Although this new design concept was applied to most of the aircraft structures, some remain designed under safe-life philosophy, such as the landing gears, as these are made from high-strength steels and are difficult to inspect for cracks in timely manner [5]. Military aviation maintained a safe-life approach verified by full-scale fatigue testing to several lifetimes, but the crash of the F-111 swing-wing fighter/bomber, on December 22, 1969, showed the limitations of this approach [6]. The failure occurred in the lower wing pivot plate, and was originated at a forging lap incorporated during the primary metal-working operation. Due to the proximity to a vertical reinforcement rib, it was not discovered in any of the production-level inspections. In 1974, the Military Specification – Airplane Damage Tolerance Requirements, MIL-A-83444 [7], was issued. This new approach differs from the fail-safe approach used after the Comet accidents in civil aviation, by including the assumption of initial damage and the possibility to have inspectable or non-inspectable structures during service life. With damage tolerant design, the concept of slow flaw growth was introduced which must characterize noninspectable structures (initial damage must not grow to a critical size causing failure during the design service life) [8]. The Dan Air Boeing 707 crash in 1977 and the Aloha Airlines Boeing 737 accident of 1988 put the emphasis on inspectability in

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Annual fatal accident rate (per million departures)

50

Rest of the world U.S. & Canadian Operators 1993 Through 2012

40 2.0

Rest of the world U.S. & Canadian Operators

1.5

30

1.0 0.5 0.0 94

20

96

98

00

02

04

06

08

10 12

Year

10

0 60 62 64 66 68 70 72 74 76 78 80 82 84 86 88 90 92 94 96 98 00 02 04 06 08 10 12

Year

3. Structural design philosophy Through the evolution discussed in the previous section, modern structural design philosophies were developed to the present day. The major design philosophy used today is mainly based on the damage tolerance approach. The different design concepts will be addressed next. Through the years, structural design concepts have evolved, going from safe-life, passing through fail-safe, to the current industry norm – damage tolerance philosophy. All design concepts have requirements in terms of input information, but damage tolerance design requires knowing potential failure modes, and associated materials characterization including fatigue and fracture behaviour as well as corrosion resistance and environmental influences, among other properties.

Number of cycles

Log Stress Amplitude

civil aviation, proving that fail safe design concepts alone were not sufficient without adequate inspections for the condition of the aircraft and its operational environment. The latter accident was mainly due to the combination of corrosion and fatigue which led to widespread fatigue damage (WFD) [9]. The corrosion damage occurred associated to the environment in which the aircraft operated (very humid environment). Although it is a low probability phenomenon, the growth and coalescence of multiple small cracks caused by fatigue and corrosion, normally referred as multiple site damage (MSD), is a dangerous possibility, as it may lead to failure of several adjacent structural elements, WFD [10]. These two accidents also showed the danger of WFD, as it may rapidly decrease the residual strength, with a loss of fail-safe capability both in terms of residual strength and adequate time for inspection [5]. The evolution of the design philosophy, as well as materials, manufacturing processes and operational safety, resulted in a decrease of fatal accidents in civil aviation, as reported in [11] (see Fig. 1).

Stress Amplitude

Fig. 1. Safety record – worldwide commercial jet fleet [11].

Ps [%] 10 50 90

Log number of cycles Fig. 2. Example of S–N curve. (a) – regular scale (b) – logarithmic scale.

3.1. Safe-life design In the safe-life approach, failure is assumed when cracks are first formed, and as such it does not consider the possibility of crack growth. Extensive fatigue tests are required for structures design under this philosophy, due to the considerable scatter these

tests present. The maximum life is then obtained by accounting the mean life and a safety factor, usually of 1.5 regarding to the ultimate load for non-detectable and acceptable damage in civil aircraft structures, obtaining a low probability of fatigue failure. Although the mean life is divided by a scatter factor, decreasing

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the probability of failure, the presence of undetected initial damage may lead to shorter fatigue lives [12]. The presence of unanticipated structural damage, which caused some known disasters, lead to the downfall of this methodology. Airplane fuselage now follows more modern design philosophies, but safe-life is still used for aircraft landing gears, and many helicopter components. Safe-life philosophy bases itself on the analysis of S–N curves (graphs plotting the magnitude of a cyclic stress – S versus number of cycles – N) and on the application of Miner's rule [13], see Fig. 2. Miner's rule is a cumulative damage model for failure presented in the following expression: k

∑

ni

i ¼ 1 Ni

¼C

ð1Þ

where ni is the number of cycles accumulated at the stress level Si , N i is the number of cycles until fatigue failure at the same stress level, and C is the consumed fatigue life, considered exhausted when C ¼1. Fatigue has an intrinsic probabilistic nature [14], not contemplated in the most efficient manner by safe-life methodology. The approach taken in this design philosophy involves using safety margins in order to compensate the scatter in the S–N curves resultant of the probabilistic nature of the fatigue phenomenon [15]. An approach taking into account the scatter in the S–N curves involves using confidence levels and probability of survival. Guidelines for calculating the confidence levels for S–N curves that may be reasonably approximated by a straight line for a specific interval of stress may be found in ASTM E0739-91 [16,17]. Another issue with the use of S–N curves concerns high cycle fatigue, as most modern technical applications, such as aircrafts, are expected to exceed 108 cycles. Although S–N curves for metallic materials have been assumed to follow a hyperbolic relationship, for some cases it has been shown that the asymptote is not really horizontal [18]. The high frequencies achieved by ultrasonic fatigue testing machines have made feasible to test materials in the range of giga cycles (109). These high cycle tests, such as the one presented in [19], indicate this decrease in fatigue strength for certain materials. This fatigue strength decrease with increasing number of cycles may be observed even if corrosion or temperature effects are not present [20], leading to failure below the conventional fatigue limit. Although the safe-life methodology is no longer generally applied in the design of primary structures of civil aircraft, many components are still designed in this fashion. In military aviation, safe-life may still be found in the design of some aircraft, as the focus on dynamic performance of these aircraft results sometimes in hard to inspect structures. The structural design methodology applied in the design of Eurofighter is discussed by Dilger et al. in [21]. The aircraft fuselage was mainly designed under a safe-life approach. The weapon system performance specification (WSPS) for this aircraft required a 6000 flight hours or 25 years inspection free service life with þ9/  3 g manoeuver envelope. The safe-life requirement was demonstrated through extensive ground testing, including full scale fatigue testing. For a safe-life designed system, it is initially required to define the spectrum, consisting of the expected external loads and environment the structure will face during its service life [22]. 3.2. Fail-safe concepts Fail-safe concepts were adopted in the late 1950s in order to address the mentioned limitations of safe-life design. Fail-safe required the use of multiple structural member concepts resulting in structures which would not fail, without leaving sufficient

opportunities for structural damage to be detected. The design envelope criteria under this methodology require analysing several failure scenarios, and demonstrating sufficient residual strength of the structure in case damage is present. Although fail-safe structures provided significantly more safe structures, some damage scenarios are not predicted, and lead to unexpected defects on in-service airplanes [23]. The fail-safe philosophy is mainly based on redundancy, which contemplates fatigue damage generally well, although redundancy does not always guarantee safe structures, as shown in the case of structures subjected to multiple site damage (MSD), e.g. the accident with the Aloha Airlines Boeing 737 [5]. The continued use of aging aircraft in civil aviation beyond typical design lives, associated with the uncertainty concerning locations for crack initiation, leads to growing concerns relating to airworthiness of the structures designed under fail-safe methodology. This uncertainty was due to limited full-scale testing and lack of teardown inspections [23]. 3.3. Damage tolerant design These concerns resulted in the FAR/AC 25.571 (Amendment 45) [24] for new airplane designs and the CAA Notice 89 [25] and AC 91-56 [26] which contemplate inspection programmes for aging aircraft. These regulation changes led to the realization that safelife, fail-safe, and damage tolerance principles all have their inadequacies and it is required to combine all the three philosophies on the development of new aeronautical structures. Military aviation has provided an important contribution for the development of this methodology, with the implementation of the Aircraft Structural Integrity Programme (ASIP) of 1975, where safe-life was replaced with a fatigue crack growth approach [22]. Damage tolerance introduced the assumption of initial structural damage, making it a requirement to consider this initial damage to determine inspection thresholds and intervals. Two qualification approaches are used in damage tolerance, slow crack growth and fail-safe [23]. Slow crack growth is applied to single load path structures as well as multiple load path structures where the residual strength capability of the structure is mainly dependent on the material's resistance to fracture [27]. Inspection intervals and their planning are in the essence of damage tolerance [23]. In order to plan the inspection intervals, it is required to couple fracture mechanics evaluations of crack growth and residual strength characteristics with damage detection assessment. Service based crack detection data in combination with residual strength and fatigue crack growth data are used to produce detection reliability ratings, accounting for multiple types of inspections and intervals in structures subjected to exploratory inspections. The introduction of damage tolerance has resulted in larger emphasis on non-destructive inspection (NDI) techniques in order to improve reliability of damage detection. The simplest NDI technique is visual inspection and is commonly employed as the first line of defence [8]. This technique uses visual light and the naked eye or some sort of magnification to detect surface anomalies. Visual inspection is very limited, as it only works for surface defects and large cracks. Dye penetrant inspection, is also very simple, and works as an aid to the visual inspection method. To perform inspections using this technique, the inspection surface is flooded with a low viscosity fluid, which is then allowed to soak for a defined time period. The low viscosity fluid will penetrate cracks, and in the end a developer is rubbed on the surface to highlight left over fluid. Some penetrants are even fluorescent, and will glow under ultra violet light. Magnetic particle inspection also relies on visual inspection, as well as electronic sensors, and may detect surface or near surface defects. In this technique magnets or

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Residual structural strength Ultimate structural strength

Strength

an electric current magnetizes the test piece, and anomalies show up as magnetic flux “leakage” [12]. In order to detect internal defects, other techniques are required. Radiography is one of the techniques that are able to detect these internal defects. This technique works by emitting X-rays, gamma rays, or neutron beams which are then absorbed by the specimen, with different densities, having different absorption rates. Radiography is especially suited to detect non-planar defects, such as voids, but it may also detect planar defects which are parallel to the test beam. Along with radiography, ultrasonic inspections allow for the detection of internal defects. This technique is a very powerful method, which allows measuring various materials, locating and quantifying defects in the surface and internally. Ultrasonic inspections work by emitting a high frequency sound beam to the test piece through a transducer and a fluid or other contact medium. Sensors receive the noise signal which has interacted with surfaces and boundaries, showing in that way the location and size of the anomalies [28]. Ultrasonic inspections have fast scanning capabilities, are low cost, allow for long-range inspection, and make possible testing inaccessible or complex components [29]. Eddy currents are another inspection method that may be used. In this technique small electric currents are transmitted to the test piece, and a reverse magnetic field is generated, which is sensed by the test coil. Due to the nature of this technique, it may only be applied in materials which are electric conductors [30]. Thermal imaging may also be used to detect defects, as some defects still show up as irregularities in the temperature field when the subject is rapidly heated and cooled. These irregularities are detected with infrared cameras [31]. Optical techniques (e.g. Moiré, Digital Image Correlation) may be used to measure in-plane and out-of-plane defects, by subjecting the test piece to a small change in the load, and analysing the strain field [32]. In damage tolerant design, the key characteristic of an NDI system is the size of the flaws that can be missed when the system is applied in the field. This characteristic is usually referred as the inspection method reliability. Due to the wide range of materials employed as well as the diversity in design of structural details, it is extremely complicated to define the reliability of each inspection technique. Due to this difficulty, in-service damage detection is aided by laboratory developed probability of detection (POD). No NDI technique presents 100% probability of detection, and in the aviation industry it is customary to characterize inspection capability in terms of the crack size for which there is a 90 per cent probability of detection, which is referred as the a90 crack size [8]. The over reliance on POD can be troublesome, especially when NDI is applied to an area in search for unknown defects instead of confirmation of known defects, as missed cracks during inspection are often not properly accounted for in POD data [23]. Another key characteristic for any NDI system to be applied in the civil aircraft for structural health assessment is the time required for inspection [33]. These techniques have evolved extensively in this regard and will continue to evolve if the amount of research in this field is taken into account, as nowadays scanning large surfaces is much more viable than it was previously, due to these improvements in testing speed. Bates et al. [33] discuss the issue of speed of inspection for composite aircraft components, and concluded that thermal non-destructive testing is up to 30 times quicker than underwater ultrasonic c-scanning. Zimdars et al. [34] also address the issue of scan speed but in this case for the terahertz imaging NDI technique, developing a highspeed large-area time domain terahertz non-destructive evaluation imaging system for CFRP structures.

5

Inspection and strength restoration

Operational spectrum

Service life Fig. 3. Scheme of damage tolerance.

All of these conventional NDI techniques cannot monitor structures continuously, requiring that the structure to be monitored be removed from service. On-board sensors allow continuous on-line structural integrity monitoring, which may be advantageous by allowing condition-based maintenance, and retirement-for-cause. Conventionally, these are passive sensors that only inform on what happened to the structure (e.g. load and strain history). Active sensors (self-excited) present some advantages when compared to the passive sensors. As these are self-excited, electric excitation, which may be difficult to accommodate in the space and weight requirements of aircraft structures, is not required. Unlike the passive sensors, the active sensors have transmitter functionality, generating elastic waves in the surrounding material and receptor functionality, where the received wave is turned into an electric signal. The information provided by active sensors is also more extensive than the one provided by the passive sensors, as the active sensors allow inquiring the structure about its health state [35]. In a schematic form, damage tolerance philosophy can be demonstrated by Fig. 3, where a virtual infinite life of a structure is assured through inspection and restoration. Maintaining airworthiness of aging aircraft is a challenge that has been addressed in the last 20 years [23]. Supplemental structural inspection programmes were developed in order to address fatigue cracking detection in airplanes designed in the fail-safe methodology. The developed programmes are in accordance with updated damage tolerance regulations to reflect the state-of-the-art in residual strength and crack growth analyses based on fracture mechanics principles. Fatigue damage involves the initiation of cracks, their propagation and eventual final unstable fracture [17]. Fracture mechanics seeks to model these behaviours and its application to aerospace structures was reviewed in this journal by Newman [36], whereas a more recent review emphasizing non-linear behaviour is presented by Schwalbe [37]. Metallic fuselage is typically riveted [38], and rivet holes are sources of potential fatigue problems, leading eventually to MSD situations, as discussed by Silva et al. [39]. The improvement of the fatigue strength of riveted joints in aluminium fuselages may be obtained using plastic expansion of rivet holes [40], a process extensively researched, including its three dimensional modelling, e.g. [40]. The field of riveted aluminium fuselage, recently the object of a book by Skorupa and Skorupa [41], has been the object of extensive research; the detrimental effect of the cold work process if not properly performed, and the effects of out of plane deformation that may be associated to its use, are issues of current interest.

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An area of interest in metallic fuselage is the replacement of riveting for welding, leading to reductions of part count, fabrication lead times and weight. Damage tolerance of such structures must be studied in detail, since further to the localized variation in mechanical properties associated with welding processes [42], and with the possible use of multi-material joints [43], such structures are monolithic or integral, instead of being differential, i.e. they provide a continuous path for crack propagation. Requirements for research of their behaviour include stress intensity factor characterization [44] including bulging effects [45], tasks that are nowadays performed with fracture mechanics numerical methods such as the virtual crack closure technique based on energy considerations as described by Krueger [46], and a variety of other numerical techniques, as extended finite elements models (X-FEM) or boundary elements [47]. Welding processes being considered include laser beam welding (LBW) and friction stir welding (FSW), and the importance of residual stress fields in such applications could not be over-emphasized, as illustrated by experimental results presented and discussed in [48–50]. The possibility of using residual stress fields to retard crack growth in aeronautical welded integral structures is an interesting development, with encouraging results presented by Schnubel et al. [51,52]. An alternative discussed by Uz et al. is the use of crenellations (systematic thickness variations) in the skin and stringers of the laser beam welded (LBW) stiffened panels, and results are promising, [53]. Another path for progress in metallic fuselage design and fabrication is the increasing use of improved alloys, such as Al–Li, a path discussed in this journal by Starke [54]. Moreira et al. show the good fatigue crack growth performance of friction stir butt welds of the current generation of 2195 Al–Li alloys [55], whereas Tavares et al. found good fatigue performance in friction stir tailor welded blanks of 2198 Al–Li [56]. The benefits of laser and shot peening in the fatigue crack growth behaviour of friction stir welded panels of 2195 Al–Li are presented by Hatamleh [57]. Aging aircraft are covered by part 26 of the Code of Federal Regulation [58], which directs Type Certificate Holders (manufacturers such as Boeing and Airbus) to support operators to develop a damage tolerance based maintenance programme for all structural repairs and alterations [59]. This is applied to all past, present and future civil aeronautical transports. These requirements lead to the development of the low stress criterion for structural repairs. This criterion defines the stress levels for structural repairs to fatigue critical baseline structures in which no damage tolerance evaluations and damage tolerance inspections are required in order to satisfy the regulation. This criterion is either based on fatigue (crack initiation) or damage tolerance (crack propagation). When based on crack initiation, it involves extensive testing including testing the recommended repair configurations, obtaining S–N curves. As different materials have different fatigue lives, like the difference between repairs of 7000 series and 2000 series aluminium's, the low stress criterion has to be defined accordingly. For 7000 series aluminium's, the “low stress level” was set at 3–4 times the design service objectives, and for the 2000 series aluminium's the level was set at more than 5 design service objectives [59]. Under damage tolerance (crack propagation), the Damage Tolerance Rating [23], which is a risk analysis metric that defines the risk of component failure at the end of design life, is obtained for the baseline using routine inspections conducted by the airlines. Keeping all the parameters constant, effective stress ratings are gradually reduced and damage tolerance rating forms are developed for each of the reduced effective stress rating. Similarly to the baseline, the damage tolerance rating is obtained for each level of effective stress rating for each repair. The effective stress rating levels are then plotted against the damage tolerance rating, and the low

stress criterion is based on the stress levels which provide many times the required damage tolerance rating. Whilst composites with a metallic component, as GLARE, are used with a view of saving weight whilst assuring high strength [60,61], most current composites applications in aeronautics are however carbon fibre composites. The Handbook for Damage Tolerant Design [8] gives in Section 1.1 an historical perspective on structural integrity in the USAF, mentioning the origins of the Aircraft Structural Integrity Programme – ASIP, conceived to ensure the required aircraft structural characteristics at USAF. According to MIL-STD-1530C (USAF), [22], the objectives of ASIP are (quoting verbatim)

 define the structural integrity requirements associated with   



meeting Operational Safety, Suitability and Effectiveness requirements; establish, evaluate, substantiate, and certify the structural integrity of aircraft structures; acquire, evaluate, and apply usage and maintenance data to ensure the continued structural integrity of operational aircraft; provide quantitative information for decisions on force structure planning, inspection, modification priorities, risk management, expected life cycle costs and related operational and support issues; and provide a basis to improve structural criteria and methods of design, evaluation, and substantiation for future aircraft systems and modifications.

This standard evolved throughout the years, mainly on the basis of experience accumulated with metallic fuselage aircraft. This provides an example of the need to update existing practice to cope with the ever increasing usage of composites in fuselage structures [62]. Specificity of composites creates a need for emphasis on impact damage criteria and climatic/environmental issues, as these are known to be potential sources of loss of integrity in such structures. The use of composites in aircraft is extensively discussed in this journal by Soutis [63] and by Kashtalyan and Soutis [64], whereas current developments in modelling and of damage mechanics is provided by Vogler et al. [65] and Camanho et al. [66].

4. Structural health monitoring The now established damage tolerance design philosophy allows aircraft to maintain airworthiness for long periods of time, by maintaining inspection and strength reestablishment programmes [23]. Although running the aircraft for these longer periods of time decreased their overall operational costs [8], inspections requiring teardowns are time consuming and costly. Efforts towards developing reliable design methodologies integrating Structural Health Monitoring (SHM) systems in aircraft seek to improve safety simultaneously with reducing inspection lead time, operation and monitoring costs. SHM encompasses several scientific and engineering disciplines, as non-linear dynamics, materials science and technology, electronics, data mining and statistical pattern recognition aiming at real-time management of the structural health of structures. Aerospace structures are commonly inspected using nondestructive techniques as visual and ultrasonic inspection, Eddy currents, acoustic emission, radiography, thermography and shearography [67]. The substitution of these costly and time consuming procedures by built-in SHM systems integrating sensors and actuators, and based on models for damage location and characterization leads to damage detection and monitoring of

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structures in real-time, thus reducing life cycle costs [68]. Sensing systems may be active or passive – whereas the first contains sensors and actuators, the second contains just sensors. Manufacturing processes for integrating a network of sensors and actuators into composite structures have been proposed, e.g. [69]. Advances in SHM sensors as piezoelectric, fibre Bragg gratings, Microelectromechanical systems (MEMSs) and strain-gauges include their applicability in increasingly harsher environments, as illustrated by the use of fibre Bragg gratings in temperature measurements during friction stir welding (FSW) of aluminium alloys [70], a joining technology of current interest for applications in metallic fuselage. Delamination failure is a major concern for composite structures. Delamination identification techniques have been proposed to quantify and locate the damage using piezoelectrics built into laminated composite structures, e.g. [69,71,72]. A drawback of piezoelectric sensors concerns the localized nature of the recovered data implying that if the sensors are not situated near the damage, its characterization may be imprecise [73]. Hybrid piezoelectric/fibre optic diagnostic uses piezoelectric actuators to input a controlled excitation in the structure and fibre optic sensors to recover the resultant signal response. Advantages of the hybrid system include minimum noise between actuator/sensor, in situ material property classification, detection of delaminations, corrosion and other damage and classification of the loading environment. In addition, hybrid systems can be used for other measurements such as temperature sensing, hydrogen sensing and acoustic emission [74]. A mathematical model for determining delaminations using an embedded layer of magnetostrictive thick film which makes it possible to move the sensing apparatus on the surface of the composite structure for damage evaluation and requires access to only one side of the structure without need for disassembly is discussed by Chen and Anjanappa [75]. The selection of diagnostic waves and signal processing methods is important for obtaining reliable damage assessment and structural integrity interpretation [76]. Integrated Vehicle Health Management (IVHM) involves tools, technologies and techniques for detection, diagnosis and prognosis of

Fig. 4. Example of 3D SMART layer [75].

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faults. Esperon-Miguez et al. [77] discuss the application of IVHM in legacy vehicles, a more challenging problem than its integration in new designs. Diagnostic and prognostic tools are usually classified as data-driven or model-based methods. The first consists of techniques to find out hidden patterns in data to determine which component or module is causing failure, or to estimate the remaining useful life of a component. Model-based or physics-based methods are based on the availability of sensors data and knowledge of the behaviour of the system under healthy and faulty conditions. The research on SHM is ever expanding, with applications found in aerospace, civil and mechanical engineering infrastructure. Farrar and Worden discus the progress of SHM in their preface to a theme issue of the Philosophical Transactions of the Royal Society-A dedicated to SHM [78]. The issue includes a general state of the art paper by those two researchers [79], recalling that the damage state of a system can be characterized answering the questions (I) (II) (III) (IV) (V)

Is there damage? What is its location? What is the type of damage? What is its severity? Given the damage and service conditions, what is the remaining useful life?

A major source of damage is fatigue, and SHM of fatigue in aircraft structures is comprehensively dealt with by Boller and Buderath in the already mentioned theme issue [80]. Lopez and Sarigul-Klijn [81] argue that statistical pattern recognition techniques perform well for clearly defined types of problems, but their effectiveness is very much limited by the quality and representation of the data from which it is trained. Data mining techniques lead to increases in data volume, and this raises needs for improved techniques for identifying the relevant data among all the data amassed. The development of inspection techniques for in situ monitoring, make it possible to inspect large areas and provide reliable and quantitative structural heath information, such as defect type, location and severity of the defect, may minimize the need for periodic teardowns as required nowadays [29]. On the long run it is expected that structural health monitoring will allow for the periodic inspection procedures to evolve to condition based maintenance. The implementation of these structural health monitoring techniques would then bring considerable savings, taking into account that structures designed and operated under damage tolerant philosophy have virtual infinite life [82,83]. One technology that has been widely studied for this application has been piezoelectric wafer transducers [29]. These small and conformal transducers are capable of being either surface mounted or embedded within structures, and generate and receive ultrasonic guided waves. As mentioned before, ultrasonic guided waves have been used for NDI of various defects on aircraft structures. Ihn and Chang [84] address the use of these sensors for the detection of fatigue cracks in metallic structures. In order to prove that the concept was viable, a physics based damage index was developed, which relates the measured signals to crack growth, and was then used in a simple fatigue test on a notched aluminium plate, showing good correlation with the crack growth measurements performed with visual inspection. Having done a proof of concept, a prototype of the diagnostic system was then constructed [85]. A riveted fuselage joint and a composite bonded repair patch with a cracked aircraft part were the two selected scenarios, as they are common sight in aircraft operations. For the first case, the proposed technique was able to measure crack lengths as small as 5 mm with certainty comparable to NDI techniques already in use, such as ultrasonic scans. Regarding

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the composite patch, results were also promising, showing good correlation with visual inspections, leading to believe that the sensors would be able to measure the crack propagation in case of debonding of the patch. Research work has resulted in the development of a commercial product utilizing an embedded network of distributed piezoelectric actuators/sensors that can be surface-mounted on metallic structures or embedded inside composite structures, named SMART Layer [86]. The ability to fabricate these sensor layers in a variety of shapes (see Fig. 4) increases their versatility and chances of possible applications for structural health monitorization. This system utilizes both passive and active sensors to perform the structural monitoring. Passive structural health monitoring (SHM) makes use of passive sensors which do not have self-excitation. Active SHM uses active sensors instead, which function more like the NDI techniques used in normal inspections, with a system exciting the structure and then reading the feedback of the excitation (e.g. eddy currents, ultrasounds). Giurgiutiu [35,87] studied the use of piezoelectric wafer active sensors for the purpose of SHM. These sensors showed that although they are small in size, and relatively inexpensive, they are able to replicate the functions of the conventional ultrasonic transducers, leading to believe in their potential for full scale structural health monitoring [87]. Tests of structural monitoring systems in real structures are of great importance in order to demonstrate its feasibility, as well as to detect possible difficulties. Zhao et al. [29] studied the implementation of a sparse array of piezoelectric sensor/actuator network in a surveillance aircraft wing section. The system was used to detect simulated defects. An algorithm for correlation analysis based probabilistic defect detection and localization was developed aiming at reconstructing the defect distribution probability map to estimate the location of the defect and to monitor its growth. Although this test was performed with simulated defects, it showed good performance in defect detection, size estimation and localization in a complex structure. Fibre optic based systems may also be employed in SHM. The sensors which show more promise are the Fibre Bragg Grating (FBG) sensors. Several fibre-optic sensor systems schemes exist, such as ring networks, double ladder bus, linear fibre lasers. A common concern with this type of sensor networks is their survivability; in order to ensure service continuity of the sensor network in the event of point failure(s) [88]. This issue is addressed in [89], where a FBG passive sensor ring is studied and network survivability and capacity for the multi-point sensor systems were enhanced. A FBG passive sensor network was developed in [90], and applied to a wing box structure. A soft computing genetic algorithm-support vector regression was developed, in order to achieve high reliability of the sensor network, and in this way compensate invalid sensors in the monitoring system. FBGs may also be applied in sensor networks with self-excitation, creating active sensor networks. In [91] a system composed of both FBG sensors and piezoelectric transducers, which are used to generate elastic waves that propagate in the structure to be inspected, was developed. The system was implemented in a subcomponent specimen simulating a carbon fibre reinforced plastic wing box structure, in a manner which would be compatible with retrofitting in an actual structure. Artificial damages were introduced in the structure, such as de-bonding and delamination, and the system proved capable of detecting the induced damages, by analysing the changes in the amplitude or delay in elastic waves.

5. Self-healing structures Alongside structural health monitoring, self-healing structures may play an important role in maintaining airworthiness of

Fig. 5. Hollow fibres self-healing [86].

aircraft structures. Self-healing structures make use of selfhealing materials, which are materials that when damaged have the ability to restore their original properties [92]. Banea et al. [93] reviewed different methods of self-healing for adhesive bonded joints. One approach at creating self-healing materials was introduced by Dry in [94], where cyanoacrylate was used as the healing agent to repair cracks in concrete. This methodology was then applied in polymer composite materials by Motuku et al. [95]. These healing systems have used either a one-part adhesive, or two-part adhesives, where the resin and the hardener and the filler may be in different fibres, or one is embedded in the matrix and the other inside hollow fibres [96]. The different alternatives for self-healing structures using hollow fibres are shown in Fig. 5. An example of self-healing structure using hollow fibres with a two-part adhesive is presented by Bleay et al. in [97]. To produce these self-healing materials, a fibre filling method was employed which involved capillary action that is assisted by vacuum. One observed difficulty in this self-healing methodology is the assurance that the healing agent reaches the damage to be repaired. In [97], the effectiveness of the repair was limited, which may be due to the adhesive curing on the mouth of the fibre. A different approach at producing self-healing materials utilizes capsules with healing agents. A polymer with self-healing characteristics which is able to repair microcracks as they appear was developed and demonstrated in [98], and made use of hollow reinforcement fibres containing a healing agent. This was achieved by incorporating a microencapsulated healing agent that is released upon crack intrusion. Upon releasing, the healing agent makes contact with an embedded catalyst, which triggers the polymerization, bonding the crack faces. Fig. 6 not only shows the rupture of the microcapsule, but also the stress state of a planar crack approaching the capsule, which was obtained with a micromechanical model. A first attempt at integrating this concept in a structural composite was performed by Kessler and White in [99]. In this study, it was concluded that the interfacial bond strength between the healing agent and the fibre reinforcement is key in the selection of the healing agent system, as it was demonstrated that debonding is the dominant mode of failure. Also, due to chosen healing agent, which produces a living polymerization, it was possible to achieve successive healing for the same specimens. A composite structure was developed making use of this concept, in order to achieve a fully self-healing structural composite system in [100]. Double cantilever beam (DCB) specimens were used to study the repair of delamination damage in a carbon fibre reinforced composite with an epoxy polymer matrix. The

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Fig. 6. Healing agent being released. a – Calculated stress state with the approaching crack in the vicinity of a sphere which is remotely loaded, b – release of the agent capture in 3 frames, c – scanning electron microscope image of healing agent capsule [87].

specimens tested revealed on average a 38% recovery of interlaminar fracture toughness. The healing temperature was also shown to play an important role, as increasing the temperature to 80 1C the healing efficiency increased to 66%. Another way of achieving self-healing structures is the incorporation of conductive components in polymer matrixes which undergo resistive healing when an electrical stimulus is applied [96]. Wang et al. in [101] identified inductive heating in carbon fibre reinforced plastic composites as a potential self-healing method. These polymers made use of a Diels-Alder based healable polymer, and microcracks were induced on them, showing the ability to self-heal through the disappearance of the same microcracks upon heating. It is also possible to apply this method of self-healing in certain metal alloys. An example of a healable system that makes use of a shape memory alloy is presented by Kirkby et al. [102]. The shape memory alloy used was a Ni/Ti/Cu composite. Wires of this alloy that bridge the fracture plane were incorporated into an epoxy composite containing wax-encapsulated Grubbs' catalyst. The alloy wires were activated through resistive heating to 80 1C for 30 min. The implementation of these wires resulted in crack closure, which reduced the total crack volume and increased the crack fill factor for a given amount of healing agent. The heating

also allowed for better cure of the healing agent, improving healing efficiency.

6. Concluding remarks A review of different structural design philosophies was performed, addressing their evolution over time, and pointing out the existing differences between the methodologies. The advances in structural design methodologies resulted in improved safety, and increased service life. The introduction of damage tolerant design made airlines integral part of the design considerations as airworthiness must be maintained through inspection programmes, and also because in-service data must be used in the design of future aircraft. Design of lightweight structures, especially civil aeronautical structures has evolved over the years, leading to increased efficiency, reliability and safety. Damage tolerance is now mandated for both new designs as well as to maintain the airworthiness of aging aircraft. Maintaining airworthiness has led to the scheduling of regular inspections which require teardowns. A review on current NDI techniques applied in these regular inspections was performed.

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Damage tolerance requires extensive knowledge of potential failure modes, materials strength to fatigue damage and environment wear, along with NDI techniques minimum detected defect and scan times. This information is collected and used to define inspection thresholds and intervals. The use of structural health monitoring within the context of damage tolerance design was discussed, and current techniques which show potential of integrating aircraft structures and provide real time structural health inspections were reviewed. SHM techniques will allow for longer periods between inspections with a high degree of confidence. Also, design of new aircraft under damage tolerant design methodology will benefit from the information collected from SHM technology, as more accurate information regarding failure modes may be extracted. Moreover, design optimization may occur due to a better understanding of the loads to which aircraft structures are subjected during loading. Selfhealing also shows potential of maintaining airworthiness of structures in an efficient manner, and several techniques to achieve self-healing materials were pointed out, revealing great potential in the future aircraft structures designs.

Acknowledgements This work was supported by the FCT project PTDC/EME-TME/ 117596/2010. Dr. Moreira acknowledges POPH – QREN-Tipologia 4.2 – Promotion of scientific employment funded by the ESF – European Commision.

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Please cite this article as: Braga DFO, et al. Advanced design for lightweight structures: Review and prospects. Progress in Aerospace Sciences (2014), http://dx.doi.org/10.1016/j.paerosci.2014.03.003i