17:00   Symposium Closure
Emiel Amsterdam, Gerrit Kool
Abstract: The use of Laser Engineered Net Shaping (LENS) techniques for fabrication and repair of metallic aerospace components has increased with increasing process development. For application of the technology on more critical aircraft structure or gas turbine engine parts, the status of the technology has to be assessed by determining the mechanical properties of relevant aerospace alloys. Process development and basic research has been done on Ti-6Al-4V and the interest is now expanding to other aerospace alloys. According to the latest process developments and parameters, two identical components have been fabricated at Optomec by the LENS™ technique. One component was made of Ti-6Al-4V and one of Inconel 718. The components were given a standard heat treatment prior to cut-up for machining the test samples. Test samples were cut from the laser track and build directions to examine the tensile properties and the microstructure for both orientations. Special attention has been paid to the gamma/gamma prime distribution (for Inconel 718), grain size, grain orientation and micro-porosity. High cycle fatigue (HCF) tests were performed on round samples to create a stress vs. cycles curve for both alloys. The fracture surfaces of the tensile and HCF samples were examined for both alloys and special attention was paid to crack initiation and its relation to the microstructure. For a good assessment of the current LENS technology status the results for Ti-6Al-4V are compared with results from (LENS-) literature, and owing to the same component geometry and process conditions the results on Inconel 718 can be compared with those of Ti-6Al-4V and forged IN718.
Eric Paroissien, Cat Tan Hoang Ngoc, Harvind Bhugaloo, David Ducher
Abstract: In the frame of aircraft performance efficiency enhancement both by reducing manufacturing costs and increasing weight saving, the design of longitudinal metallic joints of civilian aircraft fuselage is under consideration in this paper. The longitudinal junctions of fuselage are mainly single-lap bolted joints; moreover, between both joined sheets, a layer of sealant is applied in order to ensure the sealing of the pressurized cabin and the protection against galvanic corrosion. The critical criterion of design of these joints is the fatigue strength, thus justifies the need for three lines of fasteners in order to reduce the load transferred on the extreme fasteners. The hybrid (bolted / bonded) – quoted HBB – joining technology allows for associating a discrete transfer mode with a continuous transfer mode, each one having its own stiffness. The bolted system (discrete transfer mode) generates a high stress concentration around the holes, which is penalising to the fatigue strength. The bonded system (continuous transfer mode) allows a better distribution of the load transfer between both adherends. In aircraft structure assembly, the HBB technology, reduces the load transferred by the fasteners, and hence improves fatigue life, while ensuring the static strength under extreme loads by the fasteners. The target is to design the HBB joint to share the load between the adhesive and the fasteners in a suitable way. It has been experimentally shown [1], [2], [3] the possibility to obtain with HBB joints, higher static failure load and a longer fatigue life than the corresponding bonded or bolted joints with a suitable adhesive. This paper aims at showing, both by a simplified analytical approach and an accurate three-dimensional finite element analysis, that the application of the HBB joining technology instead of the classical bolted technology increases the fatigue life of longitudinal joints of fuselage. The analysis is performed by considering a three and a two fastener lines single-lap joint, representing fuselage longitudinal junctions. The adherends are in clad aluminium 2024 T3 and are linear elastic isotropic and the fasteners are in titanium and are assumed linear elastic. The analysis is composed of two steps. The first aims at defining a suitable adhesive type by using an analytical approach [4], [5]. The model employed assumes that the material behaviour of all components is linear elastic isotropic. After a brief description of the model and of the suitable development required in the frame of this analysis, it will be demonstrated that a flexible adhesive could be a suitable candidate. As it is required for the adhesive layer not to fail within the aircraft life, in the same way, as the current sealant layer, it is assumed that the flexible adhesive layer does not fail under fatigue loads. It means that fatigue life of HBB joints with flexible adhesive occurs in the adherends at fastener lines. The second is to calculate the stress concentration factor firstly using simplified approaches [6], [7] by computing the bolt load transfer provided by the analysis tool of the global mechanical behaviour. The eccentricity of the load path implies a geometrical non linearity, and in order to suitably simulate the material behaviour of the flexible adhesive non linear constitutive law is considered. Hence, secondly, to accurately simulate the mechanical behaviour of the two lines single-lap HBB joints, a non linear 3D finite element model is developed. The flexible adhesive material behaviour is simulated by a quasi incompressible hyperelastic law, under the shape of the Mooney-Rivlin potential with two parameters. The stress concentration factor is then deduced. REFERENCES [1] FU M., MALLICK P.K. Fatigue of hybrid (adhesive/bolted) joints in SRIM composites International Journal of Adhesion & Adhesives, Vol. 21 (2001), pp. 145-159 [2] KELLY G. Quasi-static strength and fatigue life of hybrid (bonded/bolted) composite single- lap joints Composites Structures, Vol. 73 (2006), pp. 119-129 [3] PAROISSIEN E. Contribution aux assemblages hybrides (boulonnés/collés) – Application aux jonctions aéronautiques Doctoral Dissertation, Institut de Génie Mécanique de Toulouse, Université Toulouse III, Toulouse, FR, 2006 [4] PAROISSIEN E., SARTOR M., HUET J. Hybrid (Bolted/Bonded) Joints Applied to Aeronautic Parts: Analytical One-Dimensional Models of a Single-Lap Joint Trends and Recent Advances in Integrated Design and Manufacturing in Mechanical Engineering II, Springer Eds., ISBN 978-1-4020-6760-0, October 2007, pp. 95-110 [5] PAROISSIEN E., SARTOR M., HUET J., LACHAUD F. Analytical two-dimensional model of a hybrid (bolted/bonded) single-lap joint Journal of Aircraft, AIAA, Vol. 44, No. 2, March-April 2007 [6] NIU M.C.-Y. Airframe structural design – Stress analysis and sizing Conmilit Press Ltd, 2nd Eds., January 1999, ISBN 962-7128-08-2 [7] SCHIJVE J. Fatigue of structures and materials Kluwer Academic Publishers, 2004, ISBN 0-306-48396-3
Riccardo Rodi, Rene Alderliesten, Rinze Benedictus
Abstract: Fibre Metal Laminates (FMLs) have been developed in the past to increase the fatigue characteristics of laminated aluminium structures by adding fibres in the bond line. The fibres are insensitive to the occurring fatigue stresses in FMLs and bridge the fatigue cracks in the metal layers by restraining the crack opening. This results in a complex mechanism of crack growth in the metal layers and delamination growth at the interface between metal and fibre layers that, if in optimal balance, results in the excellent fatigue characteristics for which FMLs are known. Next to the excellent fatigue behaviour, FMLs structures show also higher residual strength compared to an equivalent monolithic metal structure. When a high static load is applied to a FML, plastic deformation occurs in the metal constituent, as it happens in a monolithic metal. However, in the case of FMLs the fibres play a fundamental role by carrying part of the load that is “lost” in the yielded zone of the metal. Complex load transfer mechanisms take place at the metal fibre interface inducing local static delamination and eventually fibre failure. All these complex mechanisms, both for fatigue and static load cases, have been investigated in the past by many researchers. The investigations provided qualitative and, not always, quantitative descriptions of the mentioned mechanisms, like crack tip plasticity, fatigue delamination shape, bridging stress and fibre failure. The paper presents an investigation on the main fracture mechanisms occurring in FMLs for both fatigue and static load cases. Quantitative measurements of the crack tip strain field, fatigue delamination shape and bridging stress have been carried out by means of Digital Image Correlation (DIC). The tests have been performed on different kind of FMLs with different material constituents and different lay-ups (Al 2024-T3, Ti-6Al-4V and AISI 304 have been selected as metal constituents, while S2-glass, M30SC carbon and Zylon have been selected as fibres/ prepreg system). The crack tip strain field analysis provided an insight into the FML deformation behaviour and in particular the influence of the metal and fibre stiffness on the plastic zone extension and shape. Moreover, elastic-plastic FE analysis has been performed in order to study the load transfer mechanisms occurring at the metal/fibre interface due to the metal plasticization. With the digital image correlation (DIC) technique it has been also possible to visualize the fatigue delamination shape by measuring the difference in strain between the delaminated and the non-delaminated area. Measurements have been performed at different fatigue crack lengths by taking pictures under a static load equal to the maximum fatigue load cycle. Correlation of the so obtained images enables visualization of the delaminated area and finally the DIC measurements have been compared with those provided by chemical etching. Following the same measurement approach, “inverted” lay-up laminates have been manufactured in order to investigate the fibre layer behaviour. For this configuration the prepreg layers have been bonded externally providing a clear visualization of the phenomena occurring on this constituent. In particular, the bridging stress has been measured in the wake of the fatigue crack and the so obtained results have been compared with those calculated analytically. This paper presents a detailed quantitative description of the most important failure mechanisms occurring in different FMLs structures, providing a global view of the material behaviour. Moreover, quantitative relationships between the mentioned failure phenomena are provided. Digital image correlation showed to be the best technique to perform detailed non destructive in-situ strain measurements for FMLs. The understanding learned during this investigation has created a solid basis for the future development of a prediction model for residual strength in FMLs.
Mustafa-Volkan Uz, Mustafa Kocak, Francois Lemaitre, Jean-Christophe Ehrstrom, Stephan Kempa, Frederic Bron
Abstract: Laser beam welding (LBW) technology has been successfully used in production of lower body panels of A318, A340-600 and A380 aircrafts. In combination with new materials, these panels stiffened by LBW stringers offer significant advantages in terms of cost and weight savings compared to riveted (differential) panels. The strength undermatching effect of LBW, which may impose static strength limitations for loading direction transverse to weld seam is currently solved by introduction of local reinforcements (sockets) under the stringers. Damage tolerance of integrally stiffened panels can also be improved for loading case parallel to the stringers to extend the application area of these panels to more severely loaded fuselage sections. The SIF profile of a cracked stiffened panel can be modified by the introduction of thickness variations (crenellations) –to inner surface of the skin- parallel to the stringers. This geometrical modification of the skin sheet will lead to a substantial decrease of the overall fatigue crack growth rate compared to conventional panel geometry. This new design concept, ultimately leads to a “tailorable” damage tolerance behavior, which is exceptionally important for the design of the future metallic panels. This recently developed concept appears to be a very promising approach for the improvement of damage tolerance of metallic airframe components in specific locations where one or two-bay crack scenario applicable. In order to design crenellation patterns that can successfully respond to the given damage tolerance requirements, the fatigue behaviors of the crenellated panels should be predictable. In this context, this study aims to report the results of the investigations carried out to clarify the crack growth retardation mechanisms and prediction of fatigue lives of crenellated panels. In this study, superior performances of the crenellated panels over conventional “smooth” LBW stiffened panels were demonstrated by test results of large scale panels made of Al2139-T8 both under constant amplitude and spectrum loading conditions. The basic mechanisms of crack growth retardation effect of crenellations were explained and the fatigue life prediction methodologies discussed. Finally, using these methodologies, fatigue lives of large stiffened panels (stiffened with 5 laser beam stringers) were analyzed. Predictions showed good agreement with the test results. It was shown that, it is possible to calculate the retardation effect of crenellations on fatigue crack growth quantitatively with high precision. In addition to fatigue crack growth, a possible effect of crenellations on fatigue crack initiation was also investigated. For this purpose, large scale stiffened crenellated panels (740mm x 1200 mm) made of Al2139-T8 material was tested under constant amplitude and spectrum loadings. Careful examinations of panels and fracture surface examinations showed that the crack initiations did not occur at the edges of the crenellations. Aim of this paper is to report above mentioned results of a large ongoing research program on the crack growth retardation effect of crenellations on aerospace aluminum alloy panels.
Aakkula Jarkko, Saarela Olli, Kari Lumppio, Tapani Haikola
Abstract: The difficulties in developing durable and robust surface preparation techniques have limited the use of bonded joints in highly loaded metal to metal and metal to composite structural joints in aircraft applications. The PAA and -GPS silane based techniques have matured to an acceptable level with aluminium. Some difficulties are still experienced with -GPS silane on titanium. However, no acceptable method has been available for structural steel bonding. In the DIARC process the metal part is plasma treated in a vacuum chamber at room temperature. Ions with enough kinetic energy form a thin (from nanometers to microns) well adherent amorphous and dense nanostructure when they hit the surface. The process is environmentally safe and economical. The surface is dense, hard, corrosion resistant, has very low coefficient of friction, and is also bondable to epoxy. In addition to epoxy bonding, the DIARC process can be used for coating tools, sliding parts and molds for lower friction and wear resistance, and for replacing other environmentally hazardous corrosion protection coatings on metals (e.g. epoxy primers containing chromates or cadmium often used with high strength steels). For epoxy bonding the treated metal part does not need any additional primer. It also stays stable and is robust to the bonding process. The DIARC pre-treatment thus decreases the workload in the final bonding and also decreases the requirements for the surface treatment stage of the process. Testing has been performed with the AC-130 Sol-Gel and DIARC plasma coating surface treatment methods for metal bonding applications. Bare and clad 7075 aluminium, TiAl-4V titanium, AISI 304 and AISI 4130N steels were bonded with the 120°C curable FM300-2 epoxy adhesive. Wedge test specimens were tested at hot/wet (60°C/98%RH) environment and in neutral and salt hot (60°C) water immersion. The -GPS silane based AC-130 Sol-Gel treatment provided acceptable results at hot/wet conditions with bare and clad aluminium even without primer when grit blast or suitable mechanical abrasion techniques were used. The crack growth was gradual and the failure mode was over 95% cohesive. With titanium the BR 6747-1 primer improved the durability of AC-130 Sol-Gel treatment to an acceptable level. AC-130 Sol-Gel with the AISI 301 steel provided a reasonable crack growth but unacceptable failure mode at hot/wet conditions. The testing of the DIARC plasma coating with the 301 steel and titanium in epoxy bonding provided very good results at hot/wet environment. The crack growth measured in the wedge test was less than 5 mm in two days and the failure mode was 100% cohesive. The results were as good as the best wedge test results achieved in aluminium bonding. The DIARC plasma treatment increases the durability of bonded steel joints to the same level as composite to composite joints, which opens a wholly new option to design light bonded metal/composite joints and structures for space, air, naval, land, construction and industry applications. The tests will be continued. Immersion tests in neutral tap water and in salt (North Baltic Sea water with 0,5% salt content) water have been started. Preliminary results show that DIARC coating provides good results with steel and titanium also in immersion. AC-130 Sol-Gel treatment seems to provide unacceptable crack growth in hot water immersion tests. KEY WORDS: Metal adhesive bonding, Surface preparation, AC-130 Sol-Gel, DIARC plasma coating
Gregory Alan Shoales
Abstract: Safely sustaining the ever increasing numbers of aging aircraft in the United States Air Force (USAF) has brought a correspondingly increasing demand to determine the true condition of service-aged aircraft structural components. The only means available to precisely determine the damage state of a given structure is by what is commonly referred to as a “teardown inspection” of that structure. A very large number of these teardown programs have been executed in recent years. These include: A-10, AV-8B, B-2, B-52, B-727, C-5, C-17, C-130, C-141, EC 135, F-15, F 16, F-22, F-111, KC-130, P-3, T-37, T-38, VC-10 and numerous civil and general aviation aircraft. Unfortunately, there is little to no cross talk between organizations performing teardown programs. The lack of communication limits the sharing of lessons learned and best practices developed during the course of each program. Efforts in the past, such as the handbook published by The Technical Cooperation Program (TTCP) in 2005, have summarized policy decisions and other programmatic issues associated with teardown programs as they applied to the international military and civil community. However, the procedures and requirements to be considered during each task of a teardown program have not been documented in great detail in any published work. Because of their extensive experience in all aspects of teardown program planning and execution, the USAF Academy’s Center for Aircraft Structural Life Extension (CAStLE) was tasked by 77th Aeronautical System Wing, Wright Patterson AFB, Ohio to prepare just such a document. The resulting best practices handbook documents the task by task procedures for planning and executing an aircraft structural teardown program. CAStLE was further tasked, wherever possible, to capture lessons learned from past programs. As a result of this effort, the handbook titled “Procedures for Aircraft Structural Teardown Analysis” was published in early 2008. The handbook chapters address all teardown tasks from setting program goals and requirements through the analysis of the resulting teardown program data. This work details the handbook development effort and presents a brief summary of each chapter.
Sharif Ullah Khan, Rene Alderliesten, Rinze Benedictus
Abstract: Fibre metal laminates (FMLs) are hybrid materials consisting of alternating layers of uni-directional impregnated fibre lamina and thin metallic sheets adhesively bonded together. These materials have been developed primarily for aircraft structures as a substitute to high strength aluminium alloys due to their better mechanical and damage tolerance properties. For standard GLARE, aluminium 2024-T3 sheets and S2-glass fibres are bonded together with FM94 epoxy adhesive to form a laminate. FMLs are famous for their exceptional fatigue properties. These properties are due to the a complex fatigue mechanism. During fatigue loading, crack nucleation and growth occurs in the metal layers. The fibres being insensitive to the occurring of fatigue stresses remain intact and bridge the cracks in the metal layer. This crack bridging restrains the crack opening and reduces the crack-tip stress intensity factor resulting in slow crack growth rate. Secondly, due to this mechanism delamination starts to grow at the metal-fibre interface. The excellent fatigue properties of FMLs are due to a so called coupling process between the delamination and crack growth. FMLs have been extensively investigated in the past, especially the fatigue mechanisms. After developing the understanding and knowledge about FMLs at Delft, a generic and physically sound analytical fatigue crack growth prediction model has been developed. The model works quite well for the Constant Amplitude (CA) loading. Further research is done in order to enhance the capabilities of this model to predict under Variable Amplitude (VA) and Flight Spectrum Loading. To make the step towards a VA fatigue crack growth prediction model for FMLs from the CA prediction model, certain limitations are identified. These limitations are the lack of quantitative information about plastic zone formation and its size, shear lips and stress state. The size of plastic zone is vital information required to investigate the effect of overloads. Smaller or constant plastic zone sizes will reveal the presence of limited retardation/interaction effects. Shear lips and its shape reveal the effect of load variation on material properties as well as the influence of crack length. Presence of plane stress or strain needs to be studied for FMLs due to the presence of thin metallic sheets on the other hand the fibres and adhesive can influence the stress state under load variation. A research program has been executed in which experiments have been performed to quantify the above mentioned parameters. Three types of materials have been considered for the test program, namely Monolithic Aluminium, Metal Laminates (MLs) and FMLs. The correlation between the monolithic aluminium and MLs provides information on stress state and shear lips. While the effects of intact fibres and their influence combined with the effect of VA loading has been investigated by correlating the MLs and FMLs. Plastic zone sizes, crack tip plasticity and delamination shapes has been measured using Digital Image Correlation (DIC) technique. The effects of VA loading on these parameters has also been investigated. The sizes of plastic zone measured during these tests have been compared and evaluated using the Irwin's plastic zone size relation and a better correlation has been observed for FMLs than monolithic metal and MLs. This paper presents the experimental work and its analysis. Fatigue crack growth results will be presented in the form of plastic zone sizes visualized using DIC compared with the Irwin's relation. Influence of overload ratio on plastic zone sizes and delamination shapes will be discussed using quantitative data. Finally these quantitative results have been implemented in already existing CA prediction model to enhance its prediction capabilities to VA loading.
Ugo Mariani, Giuseppe Ratti, Marco Giglio, Mario Guagliano
Abstract: CS 29, under paragraph 29.571, requires that flaw tolerance capabilities are to be demonstrated for fatigue assessment. For metallic parts of helicopters, subjected to high loading frequencies, this means to design at stress levels which do not allow crack propagation since the use of the traditional LEFM crack growth approach leads to very short inspection intervals. The presence of a small surface flaw (scratch, dent, crack) significantly reduces the fatigue allowable of a mechanical part since the crack nucleation phase is either very short or even totally missing. Within this context the fatigue limit in the presence of flaws assumes a different meaning to that used in traditional safe-life concepts. It can be considered to be the threshold stress for non propagation of cracks emanating from the original flaw. Shot peening is a mechanical process which cold works the surface of a structural part by means of a propelled stream of spherical shots. It is used to improve the fatigue properties of the part by introducing on the surface and in a small layer underneath beneficial compressive stresses which retards or sometimes prevents fatigue cracking. After shot peening application a surface flaw is partially or totally embedded in the compressive stress field. When considering the traditional LEFM methodology, this compressive stress field is the main source of the closure of the crack, which therefore grows at a highly reduced rate. The same closure effect is also present when considering the threshold stresses for non propagation of cracks emanating from the original flaw, so that the detrimental effect due to the presence of the flaw is reduced. The shot peening intensity according the Almen scale is used to indicate the way to perform shot peening on a metal part. However, a given value of Almen intensity can be achieved by means of different combination of treatment parameters (shot size, material, velocity and angle) inducing different in-depth residual stress profiles. Instead flaw depth is the most important parameter for surface flaws. Hence the set of combinations shot peening parameters - flaw depth is very wide and the effect on the threshold to propagation can be very different with varying the combination chosen. In addition other parameters influence the compressive stress field so that theoretical methods available in the literature can lead to results not consistent with test experience. Within this context the present paper will show the effect of shot peening on the threshold to propagation of a semicircular surface flaw 0.36 mm deep for a typical aluminum alloy used for helicopter dynamic components (Al7475-T7351). A numerical procedure based on finite element simulation has been developed being the aim the prediction of the residual stress field induced by shot peening by varying the peening parameters. On the base of the numerically obtained results, an approach was developed to predict the fatigue strength of peened parts including notches or surface flaws. The results were compared to those experimentally obtained leading to the definition of a criterion aimed to the optimization of the shot peening parameters versus fatigue strength and threshold to crack propagation of small defects. Keywords: shot peening, flaws, flaw tolerance, compressive stresses, threshold to crack propagation
Wahyu Kuntjoro, M Suhaimi Ashari, M Yazid Ahmad, Assanah Mydin
Abstract: The Royal Malaysian Airforce (RMAF) operates fighter squadrons of F/A-18D, MiG-29, and Sukhoi-30 which were designed on Safe Life principle. RMAF has been conducting an intensive fatigue life monitoring program to the RMAF F/A-18D and is developing fatigue life monitoring (FLM) program to other fighters in its inventory. In the development of FLM capability, RMAF MiG-29 became the focus of the development and research. When the MiG-29 airplanes were purchased, no fatigue life monitoring program existed for this airplane. After the experience with the F/A-18D, it was decided to conduct the FLM based on the similar approach used for the F/A-18D. The fatigue life of MiG-29 was decided to be based on the wing-fuselage lug joint structure. As MiG-29 is a highly maneuverable airplane (similar to F/A-18D and other fighters of this category) and frequently conduct high-g maneuver, Low Cycle Fatigue (LCF) approach was adopted. FLM is conducted based on the real (actual) loading of the airplanes (rather than to the OEM spectrum loading), hence the actual stress history needs to be obtained. The stress spectra of the wing-fuselage lug joint, was derived through mapping of g-spectra to the 1-g stress level of the lug. The g-history was obtained from the accelerator installed in the airplane. With regard to the MiG-29 this data (and other flight data such as velocity and altitude) was provided by TOPAZ on-board system. The 1-g stress level was obtained by finite element modeling of the wing structure and lug joints. The loading on the lug was based on the load reactions on the lugs modeled as support points of the wing structure. The critical stress on the lug was obtained by modeling the lug for finite element analysis using the load that was obtained before. NASTRAN package was used for finite element modeling and analysis. Once the stress spectra developed, cycle counting was obtained by using rainflow procedure. LCF analysis requires fatigue characteristics of the lug material in the form of strain-cycle data. As these were not available, then laboratory tests using the lug material sample, needed to be done. Hardness test and chemical composition test were conducted for obtaining the lug material characteristic. Empirical approach was then employed for developing strain-cycle diagrams. When conducting fatigue analysis to the lug, notched effect was taken into account using Neuber theory. Mean stress effect was dealt with using Smith-Watson-Topper formula. Miner’s rule was used to calculate the fatigue damage accumulation. The fatigue life of the airplane was based on the fatigue life expended (FLE) index which is ranged from 0 to 1. FLE below one means the component is safe. FLE equals to one means the component has come to the end of its life. A fatigue life prediction software which incorporates the above concepts had been developed. The software is ready and can be used for the FLM of the RMAF MiG-29. The methodology used in the development of the fatigue life monitoring is also applicable to other Safe Life designed airplanes and becomes part of the RMAF effort in its Aircraft Structural Integrity Program (ASIP).
David Child, Nicholas Moyle, Alten Grandt
Abstract: Predicting crack growth behaviour is an important element of managing the structural integrity of aircraft fleets. It is common for fleet managers to employ software packages, such as AFGROW, to automate the prediction of crack growth behaviour. AFGROW has traditionally used empirically derived stress intensity factor solutions. However, stress intensity factor solutions generated using Finite Element Models (FEMs) are now being employed within AFGROW to provide a greater level of accuracy and flexibility. Irrespective of the move towards using FEMs, stress intensity factor solutions must still be appropriately verified in order for the results to be utilised in structural integrity management. Two independent research programs were conducted at Purdue University to review stress intensity factor solutions for cracked, pin-loaded lugs – a commonly used configuration in aircraft primary structure. The purpose of the research was to experimentally validate stress intensity factor solutions derived from FEMs. Component fatigue tests were conducted on corner, oblique and through cracked, pin loaded aluminium lugs for a range of lug geometries typical in aerospace applications. Crack propagation was measured using direct optical and marker banding techniques. Non dimensional stress intensity factor (geometry factor) solutions were calculated utilising a back-tracking method and compared with results from representative StressCheck® Finite Element Models. In the case of the through crack configuration, the comparison of experimental and StressCheck® derived geometry factors showed a close correlation and were an improvement to solutions provided in AFGROW at the time. Based on the results of this research, StressCheck® pin-loaded lug geometry factors have since been incorporated into the AFGROW software. In the case of the corner crack configuration, there was a correlation between experimental and StressCheck® results. However, this was highly dependant on the choice of pin-loading boundary conditions in the FEM. The results of this research have been utilised to define the default AFGROW software pin-loaded lug boundary condition assumptions.
Takao Okada, Kazuya Kuwayama, Shinya Fujita, Motoo Asakawa, Toshiya Nakamura, Shigeru Machida
Abstract:  Friction Stir Welding (FSW) is one of new weld process with the capability of welding high strength aluminum alloys of the 2xxx and 7xxx types. From a point of reduction of production cost and structural weight, FSW is expected to apply aircraft primary structure as alternative to rivet joint. However, the recent regulation for damage tolerance and fatigue evaluations of aircraft structure requires understanding the location of probable fracture origin and fatigue crack growth property. Especially, it is important to investigate relationship between fatigue crack growth property of FSW panel and residual stress on the panel, because fatigue crack growth property is affected by residual stress around the weld line. In this reason, many researchers have investigated the effect of residual stress on crack growth property. But their studies do not evaluate the effect of inclined angle of FSW weld line to direction of crack growth.  This presentation describes experimental results of fatigue crack growth test in order to clarify property for fatigue crack growth of FSW panel. The effects of residual stress on crack growth rate and redistribution of residual stress during crack growth were investigated. Additionally, the effect of inclined angel of weld line on crack growth rate and direction of crack growth were discussed. Sheets of 2024-T3 with 2mm thickness were joined by FSW butt joint. Specimens have FSW of inclined angles of 0 or 30 degree to the direction of applied stress. Width and length of the specimen are 400 and 1,000 mm, respectively and the starter notch is introduced at the center. The specimens are subjected to cyclic loading with R = 0.1 and stress amplitude is 50 MPa. Test frequency is 5 Hz. Crack length was measured by CCD and the scale. Residual stress field was determined by the hole drilling method. Furthermore, residual stress redistribution caused by the crack propagation was observed by strain gages located in front of notch.  The peak value of residual stress in longitudinal direction, about 300 MPa, was found on weld line. Residual stress outside of the weld line was apparently smaller than that in weld line. And the residual stress in transverse direction was smaller than that in longitudinal direction by one order of magnitude.  The da/dN-dK curve of the FSW specimen showed that crack propagation rate was accelerated by the tensile residual stress when the crack tip located around the weld line. After the crack tip ran through the weld line, crack propagation rate gradually close to that for base metal. The relationship between redistribution of longitudinal stress and crack length shows that tensile stress redistributes in front of the crack tip after the crack tip ran through the weld line. This tensile stress accelerated crack propagation rate of the crack crossing the weld line.  Comparison of crack path between the FSW specimen for 0 degree weld line and that for 30 degree shows that direction of crack growth was not affected by its inclined angle. Regardless of weld line angle, the crack grew perpendicular to the loading direction. Next, the relationship between crack length and angle of maximum principle stress in front of the notch is investigated. The direction of maximum principle stress is inclined to the weld line in case the crack tip is located near the line. If the maximum principal stress is a dominant factor for the direction of crack growth, the crack path has inclined angle.But actual crack path is mostly perpendicular to the weld line. This result indicated that the maximum principle stress is not the primary factor for the direction of crack growth. The direction of the maximum stress amplitude seems to be a primary factor. Crack propagation rate measured by inclined angle of 0 degree is almost same as that measured by the angle of 30 degree. It is supposed that this result is caused by cancel of crack opening stress for FSW specimen, because the maximum tensile residual stress, equal to 300 MPa, is apparently higher than the applied maximum stress, 55.6 MPa. Therefore da/dN-dKeff curve of base metal was evaluated using equation of the stress intensity range ratio, U, suggested by Schijve. However, da/dN-dK of FSW and da/dN-dKeff of base metal did not coincide, even if crack tip was located on weld line. It implied that crack opening stress for each weld line angle is not completely canceled by the tensile residual stress.  Now we are conducting the test with specimens which have weld line of inclined angle of 45 degree. In addition, measurement of crack opening stress is conducted to directly evaluate dKeff of the specimen. These results are presented at the symposium.
Llorenç Llopart Prieto, Georg Spenninger, Heike Wagner
Abstract: The use of cost effective infused materials like Vacuum Assisted Process (VAP) needs of splicing technique when manufacturing large fuselage components and structures. Splice configurations result in large constructions due to short widths of semi-finished products, where these are butt joined together. Only resin fills the area between these slotted semi-finished products, where cracks generate under high loading. Afterwards these cracks develop to delamination. In case of structural details as T-joint and tapering, among others, a critical behaviour under dynamic loading appears due to the fact that the matrix of the composite is highly charged leading also to delamination. At present, composite failure prediction tools do not take into account damage process leading to failure. This gap is fulfilled by tests allowing to macroscopic factor to be applied in the stress analysis. On the other hand, splice details can be dimensioned and secured via analytical and numerical procedures, only when delamination is predictable. Delamination and their susceptibility to growth are normally characterized using fracture mechanics principles and the strain energy release rate parameter (G). Compression after impact (CAI) test has become a key experiment to gather damage tolerance performance data during the design or certification phase of a new structure or material, involving composites. A correlation of parameters during impact (force-displacement), lead to the evaluation of the energy release rate under mode II loading. However, this procedure requires different stations, as impact of the specimen, damage area evaluation by means of Non Destructive Techniques (NDT), usually C-Scan, compression testing of the damaged specimen and evaluation of the results. An alternative procedure is the use of the Transverse Crack Tension (TCT) specimens. Here, a specimen containing a determined configuration of non continuous fibre layer is manufactured and tested under static or dynamic conditions until failure. The load cycles until delamination onset, by dynamic loading, are monitored by means of in-situ visual inspection. This work presents the results obtained for the energy release rate under mode II loading (GIIc) by means of both Compression After Impact and Transverse Crack Tension tests for different Prepreg and infusion materials (LRI) used at the present time in the aeronautical industry as T300/913, HTA/913, T300/914, IM7/977-2, HTA/RTM6 among others. It is shown that the results using both methods agree and therefore it can be concluded that transverse crack tension test is also suitable in order to determine GIIC in CFRP laminates. Furthermore, the use of TCT method let reconstruct the initiation and development of damages by means of delamination under static and dynamic loading, which will help on opening the floodgates to the next generation of sizing criteria and numerical tools.
Joel Schubbe
Abstract: Abstract A study has been accomplished to characterize the fatigue crack growth rates and mechanisms in thin and thick plate commercial 7050-T7451 aluminum plate in the L-S orientation. Crack growth and crack shielding with branching behavior of long, through thickness cracks is examined and compared to L-T and T-L oriented growth data. Compact tension specimens and the compliance method were used to determine crack growth rates. Constant ΔK data showed significant retardation of growth rate curves for the L-S orientation in the range of 10 to 13 MPa√m where branching and splitting parallel to the load axis are dominant growth mechanisms. Aerospace alloys are continually being developed, modified, tested and qualified for use in flight vehicles around the world. Extended service life is a prime consideration in every current design and both structural design and material selection is necessary to increase life and ensure safe flight. Every designer is striving to find that optimal mix of properties to include strength, corrosion resistance, fatigue resistance, stiffness, or other desired characteristics. The limitations of use of many available systems are the testing and qualification of these materials in the environment in which they are intended to survive. Testing, machining, cost of supply, and space requirements drive engineers to use well-known materials. Only small configurational or orientation changes are necessary to make a known material “less known” in any or all of the material selection property areas. 7050-T7451 is purported to be a superior aluminum alloy system, exceeding strength properties of 7085-series alloys and improving upon corrosion characteristics. Historically, the “banding” of properties for selection and design is used to simplify the down select process, narrowing choices to specific manufacturing materials. Testing for the extremes is common practice and helps to reduce the number of tests required the characterization of any specific alloy. Unfortunately, the most modern, tailored alloys tend to be highly anisotropic in nature and must be fully tested with loads representative of those expected in service for the expected orientation to understand the required range of properties for selection. The 7050-T7451 alloy is one of these highly anisotropic materials that need full characterization for use. Most strength properties are fully defined but if experimental crack growth rate data and fatigue properties for the L-S orientation are to be used for life estimates, they have not been tested at sufficient levels for these predictions. This study was initiated to generate an initial data set for prediction of growth rates in the L-S orientation and to characterize the morphology of the crack. In this case, it has been determined that the manner in which the crack progresses may be important to the failure mode for machined parts with multi-axial stress distributions. Specimens tested in this study were cut from 7050-T7451 aluminum plates of parent thicknesses 1.27 cm and 10.16 cm. L-T and T-L oriented specimens were tested to validate the test method against legacy data and L-S oriented specimens were tested to generate new da/dN vs. ΔK crack growth data plots and examine the influence of crack length on the mechanisms associated with the material crack morphology. Testing in an approximate range of 5 to 35 MPa√m was accomplished and at load ratios of R=0.05, 0.1, 0.3, and 0.5. It was found that at R=0.7, the specimen configuration defined for these tests, combined with the required ASTM loading levels, was incompatible for performing this testing. Crack initiation at the R=0.7 load ratio resulted in immediate splitting parallel to the load and little or no forward crack progression. Repeat tests for the other load ratios were accomplished to examine repeatability and to compare to L-T and T-L data. A MTS 810 electric servo-hydraulic 22kip test stand was used to pin load the specimens and control was accomplished using an external MTS Flex SE Test controller with MTS PC interface software. Load considerations included the load ratio, thickness to specimen depth, K-gradients, and maintaining a small scale yielding criteria. Tests were conducted in a tension-tension mode due to specimen type and Cyctest software was used to regulate the growth loads and gradients of continuous load shedding during test. Cycle by cycle feedback was provided to enable real-time load corrections. In addition, crack tip opening displacement (CTOD) data from the MTS clip gage was recorded and used to determine crack growth rates using the crack closure compliance method. Optical checks were performed to guarantee the crack length accuracy being used for compliance calculations during test. Crack lengths were measured using a Gaertner traveling optical telemicroscope with digital readout.
Hiroaki Tsutsui, Noriyoshi Hirano, Junichi Kimoto, Takahiko Akatsuka, Hirofumi Sashikuma, Nobuo Takeda, Naoyuki Tajima
Abstract: Carbon fiber reinforced plastic (CFRP) composites have been used extensively for light weight airframe structures because of their high specific strength and stiffness. Because of the serious decrease in composites caused by damage, compression after impact (CAI) strength is regarded as an important design criterion. As a damaged part should be inspected in detail by traditional non-destructive inspection (NDI) methods, such as ultrasonic C-scan and soft X-ray, the part has to be removed from operational airframe structures usually at great expense in both labor and downtime. We aim at the aircraft application of our impact damage detection (IDD) system for composite airframe structures, as shown in Fig. 1. The IDD system consists of a composite structure with installed optical fiber sensors and a monitoring measurement system. We utilize two types of measurements for impact damage detection using optical fiber sensors, as shown in Fig. 2. One of them is an optical loss measurement. When an optical fiber is subjected to damage, local bending by impact damage, such as delamination and matrix cracking, the optical intensity of the optical fiber decreases. The initiation of damage can be judged from the degree of the optical loss. The other is a strain measurement using a Fiber Bragg Grating (FBG) sensor. This sensor can measure the strain induced by impact events. The initiation of damage can be also judged from the change in strain responses, and the damage position can be detected using the difference of the arrival time of the strain to each sensor. In FY 2002, we have carried out the development and demonstration of the IDD technologies for composite structures1). 2), as shown in Fig. 3. As the result, it was confirmed that the system could detect barely visible impact damage (BVID) in composite structures. This system can be used for in-flight damage detection as well as for on-ground damage detection, and expected to provide cost reduction for inspection and associated work. To get the prospect of aircraft application of the IDD system is a target of this development from FY 2006. For the application to a practical airframe, the durability of the system should be confirmed. As the first step, several coupon level composite specimens with embedded small-diameter optical fibers were subjected to cyclic loading. The effect of the embedment on the mechanical properties of the composites and the optical intensity of the optical fibers was investigated. Then a system evaluation test using a stiffened panel (Fig. 4) with 3 blade type stringers was carried out. The panel with 1,800mm in length and 500mm in width has 13 small-diameter optical fiber sensors including 4 FBG sensors. Those sensors are embedded in the skin, stringer and interlaminar region of the skin/stringer. The panel was subjected to several impact loading before and after compressive cyclic loading test as shown in Fig.5. In the test, the durability of the damage detection function was estimated. It was confirmed that a damaged part and an impact point were detected as shown in Fig. 6. This work was conducted as a part of the project, “ Civil Aviation Fundamental Technology Program-Advanced Materials & Process Development for Next-Generation Aircraft Structures ” under the contract with RIMCOF, funded by Ministry of Economy, Trade and Industry (METI) of Japan. References 1)Tajima N., Sakurai T., Sasajima M., Takeda N. and Kishi T., “Overview of the Japanese Smart Materials Demonstrator Program and Structures System Project”, Adv. Composite Mater., 13(1), 3-15 (2004). 2)Tsutsui H., Kawamata A., Kimoto J., Isoe A., Hirose Y., Sanda T. and Takeda N., ”Impact damage detection system using small-diameter optical-fiber sensors embedded in CFRP laminate structures”, Adv. Composite Mater., 13(1), 43-55 (2004).
Shu Minakuchi, Ippei Yamauchi, Nobuo Takeda, Yasuo Hirose
Abstract: Carbon fiber reinforced plastic (CFRP) has been used for almost all modern commercial aircrafts as a primary structural material. However, the potential capability of CFRP cannot be maximized under the conventional structural design concept, consisting of skins, stringers and frames. One of the innovative structural concepts is a foam core sandwich panel structure [1]. The integral construction consists of two thin facesheets and a lightweight foam core, and can considerably reduce the weight and the number of parts compared to conventional structures. However, it has been pointed out that the crack propagation along the interface between the facesheet and the core is the critical issue. The interface crack originates from the barely visible impact damage or the fatigue shear cracks in the foam core, and seriously degrades the structural integrity. In this context, a crack arrester (Fig. 1) has been developed by Hirose et al. [2]. The crack arrester has semi-cylindrical shape and is inserted into the interface. When the crack approaches the arrester, it decreases an energy release rate at the crack tip by suppressing local deformation around the crack. The suppression of the crack propagation has been evaluated under various loading conditions, confirming that the crack arrester can dramatically improve damage tolerance of the foam core sandwich structures. In view of practical use, the arrested crack must be instantaneously detected and appropriate measures need to be taken against the damaged area in order to maintain the structural reliability. However, the crack below the facesheet is difficult to detect using conventional non-destructive inspection techniques. This study proposes an innovative crack detection technique using two fiber Bragg grating (FBG) sensors embedded at both edges of the arrester (Fig. 2). The change of the strain distribution at the arrester edges induced by suppressing the local deformation around the crack is evaluated using reflection spectra from the FBG sensors. When the interface crack gets close to the crack arrester, the spectrum from the FBG sensor at crack side splits into two peaks due to the birefringence effect of the FBG sensor [3]. On the other hand, the spectrum from the FBG sensor opposite to the crack does not change. This is because only the arrester edge at the crack side contributes to the crack suppression. Thus, the arrested crack can be detected comparing reflection spectra from the two FBG sensors. In this study, we began by evaluating the validity of the proposed technique through numerical simulation. Finite element analysis (FEA) was conducted on double cantilever beam (DCB) and end-notch flexure (ENF) foam core sandwich specimens and the changes in the reflection spectra induced by suppressing the Mode I and Mode II type cracks were predicted. As the crack approached to the arrester, the reflection spectrum from the FBG sensor at the crack side clearly split into two peaks. On the other hand, the reflection spectrum from the FBG sensor opposite to the crack hardly changed. Hence the crack could be detected by comparing reflection spectra from the two FBG sensors. It was also revealed that the facesheet bending and the in-plane core deformation played important roles in arresting the interface crack and consequently introduced the spectral changes. Finally, we conducted verification tests. The measured spectra corresponded well with the analysis results and the crack detection technique was validated. The proposed technique enables an effective application of the crack arrester and significantly improves the reliability of the foam core sandwich structures. References [1] P. C. Zahlen, M. Rinker and C. Heim, "Advanced manufacturing of large complex foam core sandwich panels," Proceedings of the 8th International Conference on Sandwich Structures (ICSS8), 606-623 (2008) [2] Y. Hirose, M. Hojo, A. Fujiyoshi and G. Matsubara, "Suppression of interfacial crack for foam core sandwich panel with crack arrester," Adv. Composite Mater. 16(1), 11-30 (2007) [3] R. Gafsi and M. A. El-Sherif, “Analysis of induced-birefringence effects on fiber Bragg gratings,” Optical Fiber Technology 6 (3), 299-323 (2000)
Nicolas Barea, Marie Masse
Abstract: The fatigue life evaluation on tension type fitting is a difficulty in aeronautical structure designing. A multitude of fatigue calculation methods and test results exist for shear joints but very few recognized ones for tension fittings. The goal of this study is to try to define a fatigue life evaluation method on tension type fitting. Finite element analysis shows the difficulty to have consistent results, because of the high influence of parameters like friction and contact, pre-load value and its precision, way to model fastener and fitting, elements type choice. This numerical approach is heavy and can be replaced by an easier and faster approach as it is often made for shear joints. More than hundred specimens in several configurations (type of fastener, geometry, 1,2 or 3 wall fittings, loading, pre-load) have been tested and also real aircraft components failures (from Airbus A380, Embraer ERJ170 and Dassault F7X partial tests or full scale tests) have been used, thus a simple method has been developed and correlated to these fatigue tests. The method steps are described succinctly here. First of all, the geometry and loading must be well known (thickness, distance to wall, radius, hole diameter, fastener geometry, fitting tension load). Then loads are determined with tension fitting simple equilibrium according to geometry, prying effect (contact stress), and assembly stiffness. To do this, the research works of J. Guillot (ref.[1]) on thread screws assembly have been used. Knowing the loads equilibrium of the fitting, maximum stresses are calculated for the different critical sites, in radius between wall and end pad, near fastener head and in fastener. Test results have permitted to define failure criterion based on the comparison of fastener and fitting stiffness. Finally tension fitting fatigue life is estimated using the maximum stress and linear SN fatigue law depending of the fitting configuration (formed or machined fitting, flexible or stiff fitting), material type and its protection, scale and stress ratio effects. This fatigue law is correlated by test. Also fatigue life of fastener can be estimated with a method based on ESDU (ref.[2]) and our test specimens. Thus the method obtained gives very satisfactory fatigue life prediction for tension fitting, with: • 100% of test results in prediction band width [N/5;5N] • 91.2% of test results in prediction band width [N/3;3N] For the fastener predictions, results are also satisfactory, fastener test results are near the fastener law based on ESDU. This test program and obtained life prediction method give a good answer for our specific tension type fitting problem and help efficiently and economically assembly design, avoiding heavy complex finite elements models. Prediction method based on finite elements models and the final method based on fitting equilibrium and tests have been compared. The finite elements model predictions are less close to test results than the others, which is easy to understand because the approach is only numerical. Trust blindly numerical approach is a bad behaviour, and we have to pay attention not to forget to come back to fatigue test to take into account real aircraft part and fatigue dispersions. The work summarized here has permitted a better understanding of tension type fitting and its several parameters, and so to find explanations to test results or aircraft’s failures. As it has been shown tension fitting life predictions can be obtained very fast with accuracy with a simple method and generic tests. REFERENCE LIST [1] Guillot Jean (1987), Assemblages par éléments Filetés. Techniques de l’ingénieur 8 – 1987, B 5 560 –B 5 561-B 5 562. [2] ESDU. (1984), Fatigue strength of external and internal steel screw threads under axial loading. ESDU Fatigue - Endurance Data, Vol 4 Stress and Strength, Vol 5 item n°84037. [3] FDT stress department (2008), Note fatigue dépliage. FDT technical report N°E0010MM-A, LATECOERE internal note.
Mathieu Fressinet
Abstract: In term of fatigue initiation, fatigue tests are commonly stopped at 107 cycles assuming that a fatigue limit exists. Indeed, under this limit the material is not damaged by fatigue and so no fatigue failure occurs. For Aluminium, it is known that such a limit does not exist and the Wöhler’s curve is extended thanks to models like the Haibach ones or thanks to statistical approaches. Unfortunately, despite these precautions, experience feedback on in service aircrafts have shown that structural elements could fail after 109 cycles. As a consequence many studies have been carried out since the beginning of the nineties to explore the domain of gigacycle fatigue (108-1010 cycles) and to characterize precisely materials behaviour. Besides, as reaching 109 cycles at 100 Hz lasts more than 16 weeks, which is not economically acceptable, many efforts have been undertaken to increase test frequencies and to set up new test processes. Fully concerned by such a phenomenon, CEAT launched its own study which is part of a project aiming at understanding the initiation and the growth of some cracks in vibratory environment, phenomena that standard tools are not able to predict satisfactorily. To accelerate the test frequency, the idea is to use hydraulic shakers usually employed to test the robustness of electronic devices in vibratory environment. The philosophy, inspired from [1] and [2], is to excite the test specimen clamped on the shaker near one of its resonance mode. The paper describes in a first part the different stages to elaborate and validate the test procedure and in a second part the first results obtained thanks to this fatigue testing method. Thus after having chosen specimen’s general design, the first task was to model it by FEM what enables to specify the different sizings necessary to get the right mode shape at the expected frequency with the right stress level. For this study, the specimen is made in 2091T3 and designed with a stress concentration area to be representative of a real structure; it also enables to reach the expected stress level at low energy. During the tests run at more than 800 Hz, special care was also taken to ensure that no consequent damping heating appears. Actually it is a well known fact that the influence of the test frequency is related to the temperature and it is assumed that its influence is minimized if the heating is not significant. To control the stress level, gauges were glued to the specimens. Unfortunately their life duration does not exceed 106 cycles and correlations had to be done between the strain given by the gauge, displacements measured by a laser and the acceleration in entry. In fact these correlations enable to control the test trough the displacement of a chosen point on the sample and so to carry out the test with the greatest accuracy even with a broken gauge. Once the test procedure was set up, tests on different samples were carried out to verify the reproducibility of such a test. Besides tests were performed near 105-106 cycles to be compared with 4 point bending tests performed with a standard fatigue test device at 5 Hz on specimens specially designed to have the same stress concentration in the studied area. In such a way the influence of test frequency was quantified. Finally after having performed tests in the gigacycle domain, this fatigue testing methodology was used to study crack initiation on pitting corrosion and to demonstrate that very short cracks could propagate even subjected to a very low stress level. [1] Methods of testing for endurance of structural elements using simulated acoustic loading, ESDU 93027 [2] Goodman Diagram Via Vibration-Based Fatigue Testing, Tommy J. George, M.H. Herman Shen, Onome Scott-Emuakpor, Theodore Nicholas, Charles J. Cross, Jeffrey Calcaterra, Journal of Engineering Materials and Technology, January 2005, Vol.129 pages 58 to 64
Tommaso Giacobbe, Fabio Sardo, Amalia Faccioli
Abstract: In order to monitor the Eurofighter Typhoon fatigue life and allow an efficient fleet management, the Structural Health Monitoring system (SHM) is currently being used to register a large quantity of flight parameters that influence life consumption, or simply characterize the usage of every single aircraft. Furthermore, synthetic fatigue parameters are also calculated by the system in order to monitor the fatigue life consumption associated to any single flight. This amount of flight and fatigue data is being collected and stored in a data base developed for the scope. As the number of monitored flights increases, data stored within the database become more and more able to represent the effective means usage of the fleet. Currently, increasing the number of aircrafts operated by the Italian Air Force, the amount of data available in the database is becoming significant to perform first studies on fleet usage. A further feature of the SHM system, in conjunction with the data base, is to allow focusing on the analysis of some specific aspects of aircraft usage, such Load Factor (Nz) occurrences, altitude cycles, airbrake usage, etc. Hence, considering also the increasing amount of data being managed by the database, it’s becoming interesting to perform some early analyses to compare “design” and “in flight recorded” data. The aim of this paper is to present a comparative analysis between design and in-flight registered data, performed on the Load Factor (Nz) coupled with Roll Rate data. During every single monitored flight, the onboard Structure Health Monitoring system (SHM) registers the number of occurrences of predefined combinations of Load Factor (Nz) and Roll Rate values. The occurrences of these values belonging to the already analyzed monitored flights have been collected, analyzed and compared to the design values. The interest on such a study resides in the possibility to verify how much the design assumptions matched the real operational usage and consequently investigate the most critical aspects of the difficult task of predicting these real usage spectra.
Mohamed Adly Attia
Abstract: Over the last two decades, some groundbreaking research concerning materials has been published. The work reported in this paper serves two purposes. First, analysis of the properties of these materials, namely: Aluminium Lithium, GLARE, composite and CenterAl. It discusses the issues and disparity in the field of materials and puts forward wing structural design concepts that are more likely to resist to fatigue. The wing consists of three main structural parts, namely: main box, leading edge and trailing edge. Weight savings of the total wing structure pours into weight reduction in the aircraft structural weight. Weight reduction is highlighted as an important reason for exchanging metal structures with material. However, reduction in maintenance cost seems to be a very important issue since materials are more likely to resist to fatigue. In response to concerns regarding fatigue problems in metal aircraft structures, two examples are discussed here, the Airbus A380 and the Boeing 787 to represent the current state of the art in materials. In particular the Airbus A380 structural weight composed of 25% composites. Meanwhile, Boeing 787 has 50% composite usage in the structure. The fact of the matter worth noting in this respect is that the CentrAl aluminium constructions are stronger than the carbon fibre reinforced plastic constructions that have been used in Boeing 787. By using CentrAl wing constructions, the weight can be reduced. It is exceptionally strong, and insensitive to fatigue. The question remains: how to realise these benefits? The second part of the work presents the materials behaviour and formability in attempt to bridge between the theory and practice. It reviews the main features of the existing materials, presenting different results that show the major differences in Aluminium Lithium, GLARE, composite and CentrAl. The work sheds light on CentrAl and its application to aircraft wing design. It analyses and summarises the present state and future tendency in the selection of material for structure design, assisting the determination of the most suitable material combination for the targeted application. In the solution, experiments are carried out through the tensile testing machine to obtain mechanical properties and yield strength. To analyse the wing structure, four very essential tests are carried out, namely: tension and compression, bending moment, torsion and buckling. A Finite element method is employed. The achieved results demonstrate how the different materials change the development process and benefit the wing structure. The performance is discussed in detail and the results are compared whenever possible with previously published results. In particular, it lays out the applicability of CentrAl to aircraft wing structure design. Attempt in this paper is to thrash out as many aspects as space limitation would allow while paving the ground for further research.
Jerzy Kaniowski, Wojciech Wronicz, Jerzy Jachimowicz, Elżbieta Szymczyk
Abstract: The paper deals with the modelling of riveted joints in aircraft structures with Finite Element Method. Presented works were carried out within Eureka project No. E!3496 called IMPERJA. The goal of the IMPERJA project is to increase the fatigue life of riveted joints, which will lead to an increase of the aircraft service life, a smaller number of inspections and lower operation costs of an aircraft. The project assumed FEM modelling of the operating aircraft’s structure at three different complexity levels, namely considering the complete structure, a structural detail and a single riveted joint. The paper presents analyses of various rivet models and calculations of a structure and a riveted joint. In the first part examples of various models, at global and local level, were presented and usefulness of them was discussed , influence of the following simplification was analysed; • neglection of rivets in a model (elements are jointed continuously) • rivet as a rigid element (MPC) • neglection of contact phenomenon • neglection of secondary bending • neglection of residual stresses after riveting process The basis of the analysis was the asymmetric butt joint model with 14 rivets. The model which took into account secondary bending and contact phenomenon was analysed as well. The method of modelling residual stresses with temperature and thermal coefficient was used. In the second part, the example of analysis of riveted joint was demonstrated for a wing of PZL M28 Skytruck aircraft. It’s is a twin-engine, high-wing, cantilever monoplane of all-metal structure with maximum take-off and landing weight 7500 kg. A submodeling technique was used there. At first, part of the wing model, based on a CAD model, was built. It includes 7 ribs and 6 bulkheads between them. Dimensions of the model eliminate stress perturbation, connected with boundary conditions, in the area near the middle rib. It was a shell model. The boundary conditions were taken on a basis of operation data. Presence of rivets wasn't taken into account. Instead of this, parts were connected continuously (nodes were merged). The Linear model of material was used. The purpose of the part of the wing model was to gain accurate boundary conditions for next model of riveted joint on the middle rib. The behaviour of whole model is correct but stress distribution around rivets is not correct. A shell model of riveted joint was build. A boundary conditions were set on a basis of result from previous analysis. Forces, instead of displacements, were used, as boundary conditions, on account of a large stiffness difference between part models (part of wing and riveted joint model). The nonlinear model of material was used. A contact effect and secondary bending were taken into account. Thanks to that, phenomena around rivets were represented considerably better. Results from this analysis could be used as boundary conditions in a detailed calculation of one or few rivets with solid elements. Such a model was consider as well. The presented method allows to analyse phenomena that appear around a rivet in a real structure, during operation. Analyses were performed with MSC PATRAN and NASTRAN software.
Milan Krkoska, Rene Alderliesten, Rinze Benedictus
Abstract: In real life, many engineering components and structures are subjected to complex variable load conditions. Sudden changes in loading are known to significantly influence the fatigue behaviour of materials, commonly known as loading “interaction” effect. In this sense, the extension of a crack during particular load cycles can be influenced by loads applied during preceding cycles, and in fact, it can influence the crack extension occurring during subsequent single or multiple load cycles. Load interaction effects may introduce serious challenges in engineering prediction calculations. Load interaction can be studied in terms of crack growth rates changes on the cycle-by-cycle level from striation spacing measurements or they can be detected as the change in crack plane. Therefore, the fractographic observation of fatigued samples is indispensable technique for fatigue failure investigations. Several models were proposed in the past, regarding striation formation and crack extension during fatigue. These models assume simplified cracking conditions, particularly the constant amplitude loading, with crack plane more or less flat and perpendicularly oriented with respect to the direction of loading. Simplified, Mode I opening conditions are considered. Resulting crack growth models are mostly symmetrical in the shape. These models are not fully able to explain the physical concept of the fatigue process under variable loading conditions, since the fractured surfaces are rather rough; especially on the micro level. Additionally, the fractured surface of the specimen that failed under variable amplitude loading contains surface features that are not present at the fractured surface of specimen that failed under constant amplitude loading. Obviously, there is a great need for an increase in understanding of fatigue processes as detected in materials and structures during variable amplitude loading; reevaluation of existing models and development of useful concept for engineering practice that can be successively employed in design and analysis. Most literature concerns interaction between overloads, and constant amplitude cycles, both usually also tension. Several authors attributed this interaction to different mechanisms. To name a few: the state of residual stresses ahead of the crack tip, different crack closure mechanisms, material changes ahead of the crack tip (strain hardening, softening), crack front irregularities and sample geometry (plain stress/strain conditions). However, less literature is available on the interaction between underloads, particularly in compression mode and constant amplitude cycles. Understanding of such interaction is indispensable for understanding of the fatigue failure. This paper will present the details about the crack growth behavior of investigated 2024-T3 aluminium alloy in terms of crack growth rates and fracture surface appearance as a function of specific loading sequences; particularly the effect of underload (compressive) cycles and constant amplitude (tensile) cycles. Test data showing the effect of individual single underloads, repeated underloads and groups of underloads on fixed and varying number of constant amplitude loads will be presented. The main focus will be given to the crack growth rates and analysis of fractured surfaces by means of electron and light microscopy.
Gregory Wilson, Rene Alderliesten, Riccardo Rodi, Rik Jan Lemmen
Abstract: A revised approach to crack bridging problems in fiber metal laminates (FMLs) improves several analysis techniques and enables crack bridging solutions to be extended to new applications. In FMLs, crack bridging, the phenomenon in which intact fiber layers carry extra load over cracks in metal layers, is responsible for slow fatigue crack growth. That bridging acts to reduce the opening of cracks has been used as a basis for analytical models of crack bridging. Such models assume that the elongation of the fibers over the length of the disbond between layers is equal to the opening of the crack flanks, and from this a compatibility relation is derived that allows the bridging stress along the length of the crack to be determined. Typically neglected in this approach is the small amount of deformation of the metal layers between the crack flank and the delamination boundary. This paper reviews a revised formulation of an FML crack bridging model that includes an exact expression for displacement of the metal layers at the delamination boundary. Though the agreement between the results of the exact model and approximate model suggests that the extra complexity is not necessary for crack growth prediction, there are several practical applications for which the exact displacement formulation is useful. By solving for the compatibility between metal and bridging material at the delamination boundary, it is possible to simultaneously solve for the bridging forces on a cracked metal sheet between two bridging layers when the two delaminations are of different sizes. This solution can be extended to an arbitrary laminate with different length cracks in each metal layer and different sized delaminations at each interface, allowing a completely generalized bridging model for fiber metal laminates to be defined. Bridging can then be solved in FMLs under bending and combined loading, for monolithic metals with reinforcing straps, and for cracks in FMLs under bonded stiffeners and doublers. Incorporating the deformation of the cracked metal layers in addition to crack flank displacement removes the limitation of current models to bridging only directly above the crack. This enables an accurate calculation of the bridging, and thus crack growth, of two cracked metal sheets laminated without reinforcing fibers. This is also directly applicable to bonded patch repairs. Extending the analysis of bridging ahead of the crack tip also improves modeling of cracks approaching stiffeners, in both FMLs and monolithic metals. The stiffener can be treated as bridging material, limiting the displacement of the cracked metal ahead of the crack. The growth of a crack towards, underneath, and beyond a stiffener can be treated entirely as a crack bridging problem. The modified crack bridging theory can also supplement experimental methods. Digital image correlation allows visual measurement of the displacement and strain in FMLs, including in the delaminated region. The displacement predicted by the exact relation described in this paper compares favorably to such measurements. Knowledge of the exact displacement field around a crack in an FML enables the adaptation of crack opening compliance techniques to experimental measurement of cracks in FMLs, which is otherwise complicated by fibers restraining crack opening. Each of these applications of the modified crack bridging theory will be discussed in this paper, and its results will be compared to results of experiments and existing analysis methods where possible.
Michel Bode, Walter Sippel, John Bakuckas
Abstract: Structural fatigue has long been recognized as a significant threat to the continued airworthiness of airplanes. This is because even small fatigue cracks can significantly reduce the strength of airplane structure. A phenomenon referred to as widespread fatigue damage (WFD) is identified as a severe degraded condition that threatens the continued airworthiness of airplanes, is theoretically inevitable, and will be reached at some point in the life of a structure. The Federal Aviation Administration (FAA) has defined WFD as the simultaneous presence of cracks at multiple structural locations that are of sufficient size and density such that the structure will no longer maintain its residual strength. A major concern of widespread fatigue damage is that fatigue cracks are initially so small that they cannot be reliably detected with existing inspection methods. The small undetectable cracks then “link up” and grow together. This growth may be very rapid and may result in catastrophic structural failure of an airplane. To address this safety concern, the FAA issued a Notice of Proposed Rulemaking (NPRM) on WFD in April 2006. The WFD NPRM proposed that design approval holders establish for certain transport category airplanes the period of time for which it is demonstrated that the maintenance program is sufficient to preclude WFD in baseline airplane structure as well as in certain repairs, alterations, and modifications (RAMs). Comments to the NPRM suggested that inclusion of RAMs in WFD assessments should be deferred until additional information is gathered. Although there is a technical possibility of a WFD-related accident involving RAMs, there are no recorded accidents attributed to WFD occurring in properly installed RAMs. Based on this, there was general agreement among industry stakeholders that there is a low risk of WFD occurring in RAMs. In addition, there are limited resources in terms of manpower, funding, and knowledge to undertake assessments to meet the requirements. Another factor is RAM structure that may be susceptible to WFD may be limited to only one airplane or a few airplanes, whereas baseline structure typically exists on each airplane. For these reasons, the FAA concurred that WFD assessments be focused on baseline structure only. However, the FAA continues to take a proactive role and is continuing to assess the need for addressing RAMs for WFD. If the assessment demonstrates that additional actions are needed to address RAMs for WFD, the FAA may consider further regulatory actions. The work presented here details a survey of RAMs on transport airplanes that was conducted over the past year and half. It is intended to provide data to get a better understanding of the risks that RAMs may pose for developing WFD and whether they need to be assessed for WFD. Existing data in the form of Airworthiness Directives and the Service Difficulty Database were examined for trends that may indicate WFD occurrence in RAMs. Details will be provided from the survey of both retired and in-service airplanes. Surveys were conducted on retired airplanes at aircraft salvage locations and on in service aircraft at the operator’s heavy maintenance locations. Additionally, retired airplanes specimens were acquired and in-depth teardown inspections were performed to look for the presence of damage indicative of WFD. A database was developed to help analyze the data for WFD risk assessments, and an example is presented. This paper also provides a comparison of results with the survey conducted by the Airworthiness Assurance Working Group in the 1990s.
Holger Sparr, Frank Schulze, Karsten Wenke, Matthias Ziegenhorn, Thomas Fleischer
Abstract: The testing equipment for fuselage panels of aircrafts at IMA GmbH Dresden is designed for fatigue tests as a preliminary stage to the full scale aircraft fatigue test. Validating the loads for application as a component test is the main task to reflect an almost real stress state compared to the complete fuselage. The natural disturbance at the boundaries of these panels is going to be minimised due to a well-adapted test rig design and the application of additional loads of well-defined value and direction. With these measures the resulting stress state in the centre range of the panel gets close to reality. The main advantage of this testing scenario lies in the much lower testing effort, both from the mechanical and economical point of view, compared to a barrel or full aircraft test and is therefore often used to assess new design or material concepts. The super-positioning of pressure, torsion and bending loads applied through hydraulic cylinders causes a complex stress state in the fuselage panel. A numerical simulation of this process by means of a finite element (FE) analysis is the most important tool to model the actual testing setup and to adjust the loads in terms of a changing panel configuration. The research project “INNOTest I” (funded by the BMBF of Germany) is a research co-operation of IMA GmbH Dresden and Fachhochschule Lausitz (Senftenberg). The main objective in this project is to extend the current testing strategy to a new set of panel configuration with open set of geometrical parameters, wall thickness profiles of the skin or new design concepts of stringer and frames. The consideration of cut-outs (windows) or new materials gets into focus as well as testing a new variety of fuselage diameters. Especially for the evaluation of the fatigue behaviour, inhomogeneous panel configurations are of greater interest since the stress states are hardly predictable and lead to qualitatively and quantitatively different results. This is valid for the already mentioned cut-outs and additionally for non-circular panel setups resulting in special bending states due to curvature changes. An experimental assessment of such setups with respect to static loads is in preparation. The FE modelling and simulation is therefore intensified in this particular area. The results coming from these analyses can be used to draw conclusion for the loading situation of future coupon tests having refinement and sub-modelling strategies of FE models in mind. In the submitted poster the results from various FE simulations will be presented and evaluated with regard to upcoming fatigue tests.
Ligieja Paletti, Calvin Rans, Rinze Benedictus
Abstract: In mechanically fastened joints, fatigue cracks have been observed, both in test and in practice, to nucleate in one of two primary locations. The first location is close to the hole edge, along the joint net-section. The second location can be found at some distance from the hole edge, causing gross-section failure. The preference for crack to nucleate at either location is determined by the prevalent load transfer mechanism. For cracks nucleating at the hole edge, bearing has been proved to be the main load transfer contribution. When the crack nucleate remote from the hole edge, friction is believed to be the primary cause. Existing models for the load transfer in mechanically fastened joints usually ignore the contribution of friction. Neglecting friction implies neglecting the related failure scenarios. Therefore, fatigue models based on these load transfer models give predictions whose validity is limited to bearing-dominant cases. Existing models for frictional load transfer describes the behavior of the friction force acting between the sheets of the joint. Those models focus only on the contacting bodies. The drawback is that the applied load is assumed to be transferred fully by friction; this assumption is wrong in a joint, where also the effect of bearing must be considered. The knowledge of the load split between the two mechanisms must then be determined. A load transfer model which includes the contribution of friction is needed in order to improve the accuracy of fatigue models. The predominant load transfer mechanism indicates the most reliable failure model for a particular joint configuration. In order to distinguish between the two main mechanisms, a parameter is introduced which identifies the prevalent load transfer mechanism for a chosen joint configuration. In this paper a model for the load transfer in a double-shear, push-fit, bolted joint is presented. By using this configuration, it can be assumed that only bearing and friction act in the joint and all the other contributions to load transfer are neglected. A key input for the model is the pressure distribution around the hole after the tightening process. The clamping pressure acting between the plates is modeled as the Hertzian pressure distribution. The frictional load is therefore calculated by using the contact mechanics equations and the Coulomb’s model of friction. An additional input is a model for the bearing load; in this study, bearing is modeled assuming a cylindrical, push-fit, rigid pin inserted in a hole without friction. A parameter a represents the splitting factor between the two load mechanisms considered. This parameter is defined as the percentage of the external applied load transferred by bearing and depends on both the clamping force and the external applied load. An iterative procedure is established for the determination of a. The output of the analytical model allows the determination of the prevalent load transfer contribution, and, therefore, the appropriate failure model. Moreover, a transition point is identified for a certain value of a. At this point the load transfer mechanism changes from friction only to a combination of friction and bearing. This transition is reflected in a difference in crack nucleation location. Experimental evidences are presented. The model presented has to be considered as a starting point for the development of a more general model. In order to describe the behavior of more generic lap joint, the complexity of the model must be increased, including multiple fastener rows and time-varying clamping pressure distribution due to bending.
Eggert Reese, Anthony Dowson, Timothy Jones
Abstract: Fatigue vulnerable areas on aircraft structures are often concentrated around fastener or open hole locations. Fatigue enhancement methods are often applied to these critical locations prior to joint assembly to retard the potential for in-service crack initiation and growth. Cold expansion of holes (‘cold-working’) is a fatigue enhancement technique that is used extensively on metallic aircraft components to enhance the fatigue performance of a structural assembly. In addition, the cold expansion process is also incorporated into the structural design philosophy as a means of reducing the weight and cost of the structure. Cold expansion is undertaken by expanding a hole radially to an extent whereby the surround of the hole is plastically deformed. The cold expansion results in a zone of high residual compressive hoop stress in the bulk material surrounding the plastified hole. During aircraft service, the resultant residual compressive zone counteracts and reduces any tensile fatigue stresses that are normally associated with crack initiation and growth thereby improving the overall fatigue performance of the structure. Airbus currently employ two cold expansion processes on metallic aircraft structures. Although these processes provide reasonable results in fatigue enhancement of Aluminium alloys, they exhibit disadvantages like e.g. inhomogeneous strain distribution around the hole or risk of crack formation near ridges – in particular with new generation Aluminium alloys. Therefore, a new process, named the “Variable in-situ Expansion (VarEx)”-process, was developed at EADS Innovation Works. The tool employed allows the application of variable expansion levels; it provides a homogeneous strain distribution around the hole and allows to even cold work difficult to expand Aluminium alloys like AA7xxx series up to high levels of applied expansion without crack formation yielding an even better fatigue enhancement. The paper will also address additional benefits like: a precise control of the expansion level, an ability to compensate for hole tolerances and wear entailed changes in tooling diameter, a simplified process control, a improved material exploitation at a reduced risk of crack formation, a potential for joining dissimilar materials (e.g. hybrid stacks), at significantly reduced costs.
Petr Augustin
Abstract: Application of new production methods such as high speed machining, laser beam welding or friction stir welding in the aerospace industry leads to the reduction of manufacturing costs. Structures designed for these production techniques are typical by the integral character. Comparison of their behavior with differentially stiffened designs shows lack of damage tolerance characteristics and that is why this topic is investigated. The paper describes methodology of numerical simulation of fatigue crack growth under constant amplitude and spectrum loading and its application on integrally stiffened panels manufactured using high speed machining. Relatively large experimental program comprising both the constant amplitude and spectrum tests on integral panels and CCT specimens was undertaken at the Institute of Aerospace Engineering laboratory in order to acquire crack growth rate data and enable verification of analyses. The work was performed within the scope of the 6th Framework Programme project DaToN - Innovative Fatigue and Damage Tolerance Methods for the Application of New Structural Concepts. Presented methodology of crack growth simulation starts by the calculation of stress intensity factor function from finite element results obtained using MSC.Patran/Nastran. Prediction of crack propagation in stiffened panel requires determination of stress intensity factors for large number of crack configurations and that is why simple FE models comprising shell elements were built. The crack closure technique was used for calculation of stress intensity factor values. Subsequent crack growth analysis is done in NASGRO and uses description of crack growth rates by the Forman-Newman-de Koning relationship. Two crack growth models were applied for spectrum loading: non interaction and Willenborg model. A typical feature of integrally stiffened panels is branching of cracks growing through the stiffener. Simulation of this phenomenon requires consideration of parallel propagation of two cracks during the cycle-by-cycle computational procedure. This problem was treated using two-dimensional growth option in NASGRO. The methodology described above was applied for analysis of two-stringer panel tested within the experimental program of the DaToN project. The panel was made of 2024-T351 aluminium alloy using high speed cutting technique. Fatigue cracks in the skin were initiated from the central saw cut with the size of 2a = 20 mm. First analyses and verification tests of panels were performed under the constant amplitude loading at maximum nominal stress of 80 MPa with stress ratio R = 0,1 and at maximum nominal stress of 110 MPa with R = 0,5 respectively. For predictions of crack growth and experimental verifications using the spectrum loading, a load sequence representing service loading of the transport airplane wing was prepared. Applied load spectrum was measured on B737 airplane within the joint FAA/NASA collection program. The load sequence is composed of 10 flight types with different severity, analogous to the standardized load sequence TWIST. Generation of random sequence of loads and flights is realized by the in-house computer program. Before application on the stiffened panels, a calculation of crack growth under the spectrum loading was performed for simple CCT specimen geometry. Since analytical solution of stress intensity factor function is known in this case, it was possible to verify crack growth models used without the influence of inaccuracy of stress intensity factor determination based on FEM. The paper finally presents comparison of simulations of fatigue crack propagation in two-stringer stiffened panel under the spectrum loading with verification tests of two panels performed in the Institute of Aerospace Engineering lab.
Shehzad Khan, Andrey Vishnevsky, Jorn Mosler
Abstract: An Investigation on Low Cycle Lifetime of Al2024 Alloy Shehzad Saleem Khan, Andrey Vishnevsky, Dirk Steglich, Jörn Mosler GKSS Research Centre GmbH, Max-Planck Straβe 1, 21502 Geesthacht The contradictive demands of weight reduction in the constructions on one hand and extension of lifetime on the other hand lead to the necessity of using predictive tools for lifetime estimations. In this context, the choice of a particular mechanical model plays a crucial role. The presented work deals with both, experimental investigations and modeling of low cycle fatigue (LCF) damage evolution. Based on the microstructural characterization, significant work has been done in the past years by collaborators in GKSS regarding the understanding of deformation and failure behavior of high strength aluminium alloys. Taking the task to the next level the phenomenon of low cyclic fatigue is analysed here. The alloy Al 2024 T351 (thick hot-rolled plate) widely used in the aircraft industry (internal structures including extrusions and plates) is under consideration. This alloy is renowned for its good mechanical properties such as the strength to weight ratio and an improved corrosion resistance. Following the previous investigations [1] (on monotonous loading) the experimental program was defined. It consists studying of cylindrical tension-compression specimens with toroidal notches under room temperature with the frequency of loading 0.01Hz and varied strain amplitude. To capture the influence of multiaxiality of strain and stress state on damage evolution, the notch radii have been varied. To stay in the realisable range of time for the numerical simulation, the loading program, which mainly concerns the study of ultra LCF (10 …100 cycles), was proposed. In the low cycle fatigue regime, irreversible microstructural changes take places, with increasing order of strain amplitude in the form of: persistent slip bands, rearrangement of dislocation systems into cell structures, and void nucleation and growth at the secondary phase inclusions [2, 3]. The latter mechanism is peculiar of high strain amplitudes, for which very short lives are usually expected. For better understanding the mechanisms of LCF in Al 2024 alloy numerous microscopic investigations were conducted. They include the microstructural analysis of the fracture surface, the study of regions coincided to the fracture surface in micro-cuts and those on the free-surface of the specimen with the optical and scanning electron microscopy. The micro-structure is characterised by layered domains, which are bordered by intermetallic particle clusters and /or grain boundaries. It is obvious that for the monotonous loading or for the loading at extremely low number of cycles the damage occurs due to fracture of brittle intermetallic particles, whereas for higher number of cycles the microscopic analysis of fracture shows pronounced fatigue character. Moreover, specimens with smaller notch radii depict different damage mechanisms near the notch root and in the middle of the specimen. Consequently, for the modelling of LCF it is essential to describe the damage mechanisms and the resulting crack initiation correctly. An assessment of specimens’ lifetime is conducted using finite element analyses by constitutive equations including damage variables, e.g. [4, 5]. Different notch geometries and loading ranges are investigated to check the applicability of the models. The results of the simulations were compared to the macro behavior of experimentally tested specimens. The evolution of the damage variable is evaluated in the view of microstructure changes. Later the verification of the simulated results is presented. References [1] Steglich D. et al., Anisotropic ductile fracture of Al2024 alloys, Eng Fract Mech, Volume 75, Issue 12. Pages 3692-3706. 2008. [2] Klesnil, M. and Lukas, P. Fatigue of Metallic Materials, Elsevier Science Publishers, Amsterdam (1980). [3] Polak, J. Cyclic plasticity and low cycle fatigue life of metals, Elsevier Science Publishers, Amsterdam (1991). [4] Steglich, D.; Pirondi, A.; Bonora, N.; Brocks, W.: Micromechanical Modelling of Cyclic Plasticity incorporating Damage. In: International Journal of Solids and Structures. Vol. 42 (2005) 2, 337 – 351. [5] Demorat, R.: Fast estimation of localized plasticity and damage by energetic methods. Int Jal of Solids & Structures. Vol 39. Pages 3289-3310. 2002.
Pierre Madelpech
Abstract: Our general objective is to progress in the certification of metallic structure repair process by bonding a composite patch on the damaged part. On military aircraft, some countries already use that repair and its mechanical efficiency is now well established. The composite patch acts as an efficient stress bypass avoiding any supplementary hole drilling and associated stress concentration. Nevertheless the process is not used in France because of a lack of confidence in bonding. This is still today a major certification concern, with a lot of remaining needs to assure the bond quality and durability. Debonding can initiate due to wrong bonding process, to environmental aggression or as a result of a static or fatigue loading. Although this is a real problem observed on in-service patched structures and during tests, until now, no hazardous consequences have been noticed. Moreover a disbond can be detected by classical NDI. Therefore, disbond propagation can be controlled if a structural inspection program is defined and followed. With such a procedure, damage tolerance philosophy requirements may be fulfilled. The study is part of a wider damage tolerance approach of bonding defect initiating and propagating between the metal and the composite. For this purpose, initiating and propagating must be predictable with sufficient confidence. In fact, this paper only addresses the problem of predicting the growth of an initiated disbond between the metallic structure and the repair. Added to the current problem of studying bonding, the composite repair of metallic structures is more complex, since the assembly needs a curing at 120°C. Indeed such a process raises a problem, as the difference of thermal dilatation coefficients between the two materials creates residual stresses after cooling. The modelling of the disbond growth has been realised, based on total energy release rate. The numerical calculation of an aluminium plate covered by a composite patch has been performed using Samcef. The adhesive has been modelled by interface elements. It gives precisely access to the stress in the adhesive avoiding all at once too refined mesh and degenerated elements. It is essential for the study of a phenomenon occurring in the adhesive. This approach based on energy release rate does not differentiate the different kind of solicitation (related to stress intensity factors KI and KII). The initial bonding defect has been supposed to be circular and the propagation path has been imposed according to common sense but also to test results. Step by step, the disbond is modelled and the energy release rate is computed. From the propagation law, the number of load cycles necessary to cause that size increment is determined. Thus, the propagation is predicted and the life time of the bonding is obtained. The study also includes test to determine the unidirectional propagation law of a disbond. The total energy rate is more efficient when the ratio of solicitation mode (KII/KI) remains homogeneous between computing and testing. Fatigue testing was performed in order to characterize the linear elastic fracture parameters of the adhesive. The design of the fatigue coupons was therefore chosen as close as possible to repair conditions on aircraft. Moreover, it was not relevant to perform traditional tests on the adhesive as debonding can be cohesive or adhesive. The two different interfaces (Carbon – adhesive and Aluminium – adhesive) must consequently be present in the specimen. During the test, the growth has been regularly measured by common NDI and the coefficients of the Paris law have been deduced. The successive forms of the disbond have been used to define and control the modelling of the disbond evolution. Finally, it gives a tool to estimate the propagation law of a disbond occurring under a composite patch.
Jos Sinke, Sten Johansson
Abstract: In the DAILFAST program (EU KP6 program), a number of new metal and hybrid technologies have been investigated. Main objectives were to improve the properties of structural materials and concepts and to reduce the manufacturing costs. Part of this research was aimed at the improvement of the fatigue and damage tolerance properties of selected Metal Laminates and structures made of these laminates. The constituents of these Metal Laminates were state-of–art aluminum alloys, having several thicknesses, and adhesives, some of them selected for their potential for low cost curing of Metal Laminates structures. In the program a number of static properties have been tested on coupon level. The results of most tensile dominated tests showed that, as expected, the static properties of the Metal Laminates are equal to the properties of monolithic materials, since the adhesive does not contribute to the static properties of the laminates. This is true for properties like static strength, blunt notch and bearing strength for pin loaded holes. For some other properties that are related to shear and/or compression loading some improvement is achieved. However, this is not concluded on the results of coupon tests, but indirectly on the test results of some large test panels. Other properties related to fatigue and damage tolerance showed the potential for improvement. One of these properties is the impact resistance. Coupons tested in a drop tower test set-up showed that most laminates perform significantly better than the reference materials. The impact energy to failure, full penetration of the material, is 2-3 times higher than for the currently used aluminum alloys. In addition the layered structure of the laminates result in a different failure mode, this is also in favor of the laminates. Also the fatigue resistance of the metal laminates is higher when compared to monolithic alloys: the increase of fatigue life is in the order of 10-20%. These results are based on tests with specimen with through the thickness holes. Further improvement of the fatigue and damage tolerance behavior of Metal Laminates structures is achieved by the proper design and manufacture of those structures. Metal Laminates, when applied in aircraft structures, can be tailored for its purpose by selecting the right combinations of constituents and by the specification of the right local thickness. This tailoring is achieved by the application of doublers, local reinforcements and other local features which are bonded onto the skin in one bonding cycle. To demonstrate the effects of tailoring, a number of small and medium size panels have been designed, manufactured and tested within this DIALFAST program. At coupon level a number of joint types and joint configuration were tested in a static test and in fatigue. These coupon tests showed that the best joining method for metal laminates is adhesive bonding; riveted and bolted joints will not improve when monolithic materials are replaced by metal laminates. The panel tests showed that a weight reduction of 10 to 13% is possible when compared to riveted structures. The manufacture of these panels however, will be more expensive.
Elzbieta Gadalinska, Jerzy Kaniowski
Abstract: The work was done as a part of the IMPERJA Eureka Project. The goal of the IMPERJA project is to increase the fatigue life of riveted joints, which will lead to an increase of the aircraft service life, a smaller number of inspections and lower operation costs of an aircraft. The consortium intends to meet this goal by investigating and improving the riveting process as well as improving the prediction methods for fatigue life. Riveting is the most commonly used method of joining sheet metal components of the aircraft structure. Typically, the number of rivets ranges from several thousands to some millions in a single aircraft depending on the specific aircraft type and size. The riveted joints are critical areas of the aircraft structure due to severe stress concentrations and effects such as fretting and secondary bending. Therefore the fatigue crack initiation will start at the rivets holes. Fatigue crack initiation usually occurs at a number of rivet holes (multiple site fatigue), which may lead to widespread fatigue damage and reduced residual strength. Although the literature on the fatigue behaviour of riveted joints is quite abundant, many aspects are still not sufficiently understood and investigated and, therefore, they require a further study. The aim of researches presented in the paper is to determine the residual stress field, which is created during the riveting process with press (with various forces) or manually. The residual stress field around the rivets has the significant influence on fatigue life of the structure. As for now the residual stress field was determined with FEM method and it is in need to verify those results with various experiments because there are very few experimental verifications. The problems of x-ray diffractometry in stress measurements were discussed: • the problem of surface layer – cladding and anodized layer – the problem of penetration depth • the problem of measuring the gradients of residual stress • the problem of removing the rivet head and the formed rivet head to enable the measurement as close the rivet hole Measurements were performed on six types of specimens: • the specimens for determining the influence of removing the anodized layer and cladding • the specimens for determining the influence of removing the rivet head and formed rivet head. • the specimens 50x25mm – measurements before and after the removing the cladding and anodized layer. • the specimens 18x18mm – manually and press riveted - countersunk head with compensator –measurements before and after the removing the cladding and anodized layer, and after the removing the formed rivet head. • the specimens 250x60mm – manually and press riveted - protruding head with compensator – measurements before and after the removing the cladding and anodized layer, and after removing the rivet head and formed rivet head. • the specimens 681x200mm
Sandeep Shah, Scott Fawaz
Abstract: AA7079-T6 was widely used in the 1960s for manufacturing aircraft due to its high mechanical strength. Subsequently, this alloy was discontinued due to very poor fracture toughness, crack growth resistance, and corrosion resistance. A large strategic cargo carrier has aft upper fuselage (crown) skin made of this alloy. Various cracks had been identified in this area which has led to significant inspection of the aircraft. Visual inspection of the complete crown skin is done at least three times a year to detect cracks. A detailed failure analysis of several cracks from a few replaced crown skins was conducted to understand the root cause of these cracks. The failure analysis revealed that initial propagation was along the thickness of the skin which then diverged and propagated along mid thickness parallel to the surface of the skin. This crack propagation was accompanied by branching along the grain boundaries in the propagation path. Such crack morphologies are typical of stress corrosion cracking (SCC). However, study of the fracture surfaces revealed the presence of fatigue striations. Comprehensive analysis of the loading condition and operating environment does confirm that fatigue and corrosion occurs separately for these cracks. The fatigue crack propagation occurs while the aircraft is in flight and corrosion occurs while the aircraft is on ground. This corrosion weakens the material preferentially along the grain boundary leading to intergranular fatigue crack growth in flight. For rolled thin sheet product, this inter granular crack appears as an inter-lamellar crack as the grains are oriented in one line. This hypothesis was also verified with controlled corrosion-fatigue experiments in the laboratory which replicated the failure surface obtained in service. Experiments were also conducted to determine the threshold for crack growth in fatigue as well as SCC. The fatigue crack growth threshold was observed at 1.7 ksi.in1/2 and the SCC threshold was found to be 7.8 ksi.in1/2. The crown skin area where the cracks are located is loaded only during flight when the environment is benign for SCC to occur. On the ground where the environment is amenable to corrosion, there is very little load on the crown skin area where cracks are located. This results in very low stress intensity value and no crack propagation occurs by SCC. The cracks are therefore believed to have propagated during flight by fatigue loading, which is also supported by failure analysis results.
Sarah Galyon, Saravanan Arunachalam, James Greer, Matthew Hammond, Scott Fawaz
Abstract: In complex aircraft structure, crack growth rarely propagates in the idealized fashion assumed in durability and damage tolerance analyses (DADTA). Usually the applied loading is not perpendicular to the crack nucleating feature and subsequent crack propagation. This situation is known as mixed mode crack growth or in more general terms, three dimensional crack growth. Most DADTA’s conducted assume mode I only; thus, engineering judgment is used to estimate the amount of error present in the idealized models. The Center for Aircraft Structural Life Extension (CAStLE) at the United States Air Force (USAF) Academy completed a project to generate three dimensional (3D) crack growth data and predict the measured crack growth rate, crack trajectory and residual strength using state-of-the-art stress analysis and life prediction tools. Specifically, we generated the 3D fatigue crack growth test data using 1.6 mm (0.063 inch) thick aluminum alloy (AA) 2024-T351 ARCAN specimens in an ARCAN test fixture. The ARCAN test fixture allows the ARCAN sample to be rotated to produce different mixtures of mode I and mode II loading (with 0° being pure tension/compression (mode I) and 90° pure shear (mode II)). ARCAN specimens were tested at angles of 0°, 30°, 60° and 75° under constant amplitude loading and a stress ratio (R) of 0.1. The stress amplitude was adjusted to control the crack growth rate and plastic zone size. A grid on the surface of the sample was used to optically track crack trajectory and crack growth rate. While mechanical testing was being completed, we also developed a crack prediction model of the ARCAN specimen using FRANC3D/NG, a solid modeling, mesh generation and fracture mechanics code from Cornell University Fracture Group. FRANC3D/NG should be able to predict the cracking behavior observed in the ARCAN tests. A parallel effort was also undertaken to develop an engineering model of mixed mode crack growth where contributions to mode I and mode II growth were accounted for using KI and KII and the appropriate baseline crack growth data. For both the FRANC3D/NG and engineering model analysis, crack growth rate data is required and was produced per ATSM E647 using a 15.24 cm (6 inch) wide, 1.6 mm thick AA 2024-T351 M(T) specimen in both the LT and TL orientations under constant amplitude loading and an R of 0.1. The fatigue crack growth trajectory prediction for the ARCAN specimens was assessed with three different measures. The first was the point-wise comparison of the measured and predicted crack angle for each of the test conditions. The second was a block comparison of the measured and predicted crack angle. A block is defined as a discrete amount of crack growth, so the crack growth was compared at ¼, ½ and ¾ the total crack length. The third was examining the effect of cracking angle on the residual strength or critical crack size of the specimen. Comparisons were also made of the predicted and observed fatigue life, number of cycles at the end of the test and crack growth rate. The success of the prediction model is based on how the correlation of the model affects the ability to predict the measured fatigue and fracture performance based on the mechanical testing.
Hideki Soejima, Noritsugu Nakamura, Toshimichi Ogisu, Yoji Okabe, Nobuo Takeda, Yasuhiro Koshioka
Abstract: Carbon fiber reinforced plastic (CFRP) became a common material for not only secondary structures but also primary structures of aircrafts in recent years. Because CFRP has a high performance such as high specific strength ratio and high specific modulus ratio that can contribute to reduce structure weight significantly and improve fuel efficiency of aircrafts. However, impact energies and overloads to CFRP structures generate a delamination of inter layers or de-bonding of adhesive layers, which might cause catastrophic failures of the structures. Such damages are hardly visible from outside of structures, also, it is difficult to evaluate precisely such damages by visual inspection. Therefore, CFRP is applied to aircrafts with a lot of margin to assure airworthiness of aircrafts. Structure health monitoring (SHM) technologies have been investigated all over the world for several decades. In the concept of SHM technologies, sensors, which are installed in aircraft structures permanently, can detect generating damages in the structures and evaluation devices, which will be installed into main computer of aircrafts in the future, can evaluate the structure integrity by analysis of the information acquired by the sensors. The SHM systems are expected to be able to reduce life cycle cost of aircrafts and improve fuel efficiency and safety and reliability of aircrafts. Therefore, several types of sensors and principles for detection are investigated in this field. In this background, we have developed the novel SHM system that can evaluate de-bonding of adhesive layers and a delamination in CFRP structures, such as skin/stringers of wing structures. The SHM system consists of fiber Bragg grating (FBG) optical fiber sensors and PZT transducers as sensors and actuators respectively. The principle of technologies is as follows. A PZT actuator generates a lamb wave and the lamb wave propagates into a structure. Then, a waveform of the lamb wave changes with a change of adhesive condition, i.e. a change of de-bonding area. Here a FBG optical fiber sensor can detect the change of the waveform and the SHM system can evaluate the de-bonding area of the adhesive structure by analysis of the change of detected waveforms. In this study, we conducted sub-component test in order to demonstrate the evaluating capability of the SHM system, which could assess de-bonding area of adhesive structure. The sub-component test article which simulated CFRP skin/stringer bonding structure had three stringers and 3 ribs with practical pitches. Stringers were bonded to the skin by secondary bonding process. FBG optical fiber sensors were bonded to the surface of stringers by adhesive and integrated into adhesive lines between stringers and skin. PZT actuators were bonded to the same surface of the skin to which the stringers were bonded. De-bonding and delamination were introduced to the bonded sections of skin/stringer by impact using drop weight type impactor. Some impacts were introduced to the same point, and then the impact energy levels were changed gradually from low level to high level in order to simulate the damage growth caused by aerodynamic loads of aircraft. Here, we measured lamb waves by the developed SHM system before and after each impact to acquire waveforms that corresponded to each de-bonding conditions. Damage index, which was used to evaluate the de-bonding, was derived from the detected waveforms in the SHM system. In order to assess the evaluating capability of the SHM system, damage index was compared to the results of conventional non-destructive inspection method (A-scan). As a result of the sub-component test, it was confirmed that relationship between the damage index and the de-bonding area showed a reasonable correlation. As the results, it was verified that this SHM system could evaluate the de-bonding area by the damage index.
Krzysztof Dragan
Abstract: Composite structures become more popular in the aerospace industry because of their unique features and such as: excellent strength/weight ratio, corrosion resistance, possibility for complicate shape manufacturing. Extensive use of composites in Aerospace components nowadays become real and one of the examples such trend is Boeing 787 Dreamliner aircraft, where nearly 50% of aircraft structure has been made from composite elements. Increased usage of composites for aircraft structure influence on necessity for gathering information about structural integrity of such components. The reason for that is the following, during the manufacturing of composites as well as during in service and maintenance procedures there is a possibility for damage or failure occurrence. There is a large number of failure modes which can happen in such structures. These failure modes affect structural integrity and durability. In this work approach for detection of composites damage detection such as: delaminations, disbonds, foreign object inclusion and core damage has been presented. This detection is possible with the use of advanced P-C aided Non Destructive Testing NDT methods. In this work methods for detection of such damages were elaborated. Moreover, software based on Image Processing and Mathematica Library was developed to evaluate results received from automated systems. This software enables automated determining of damage size, based on numerical calculation. In this work detectability for selected methods and accuracy of elaborated software was determined. This work was done for the needs of composite main rotor blades of helicopters. These blades are made from Glass Fibre Reinforced elements GFR. In the blades structure spar is reinforced with the use Glass Roving material and honeycomb elements are made from Nomex or Glass Epoxy. This work presents modern approach for damage detection in the composite main rotor blades of helicopters. In the article will be delivered information about used NDT techniques, results of inspection and data evaluation. Moreover the most accurate techniques were described for detection of selected failure modes.
Cees van Hengel, Peter Kortbeek
Abstract: Since their conception in the late 70's of the last century, Fibre Metal Laminates (FML's) have bridged the gap between theory and practice several times, beginning from ARALL on the C17 aft cargo door, and more recently as (standard) Glare and as HSS Glare on the A380 fuselage. This paper summarizes these evolutionary cycles, and outlines some of the activities currently in progress in preparation for the next cycle. Although it is not always remembered today, the first FML generation to successfully bridge the gap was ARALL. ARALL, for Aramide reinforced aluminium laminate, was conceived of in the late 1970's and early 1980's as a highly desireable "fatigue insensitive aluminium" material The theory was to find the optimum combination of the newly developed fatigue insensitive high strength Aramid fibres with the time proven production technique of adhesive metal bonding. Major development efforts followed, such as for lower wing skins on the Fokker 50 commuter airplane. Ultimately "the gap" was bridged when ARALL entered the flight demonstration phase when it was applied as the skin material for the aft cargo door on the McDonnell Douglas C17 military airlifter. So, FML 's had completed their first evolutionary loop: starting from theory, ARALL had gone through material specification, material qualification and allowables development, design and manufacturing, and finally reaching operational useage. However, as soon turned out, "survival of the fittest" for aerospace materials demanded more than "just" good performance on properties: good performance on cost was also imperative. ARALL was manufactured as flat sheet laminates, pre-stretched to counter unfavorable curing stresses. The double-curved skin part was manufactured using conventional metal processes: individually forming the sheets by rolling or stretching, cutting and assembling with titanium straps. The resulting significant costs were not justified by the weight saving, and so, after a limited number of ship sets, ARALL was again replaced by conventional aluminium. Some 10 years later, the second generation of FML's have managed to bridge the gap again, this time as Glare on the A380. The original idea behind this FML generation had "just" been to make an FML with biaxial reinforcement, which required the use of glass fibres to avoid the need for prestretching. As expected from the theory, the new material did indeed show good fatigue properties, but also unexpectedly performed well under impact, lightning strike and fire (burn through). Nonetheless, as with ARALL before, cost constraints seriously threatened Glare's establishment as a viable aircraft material. This time however, the new fibre ingredient allowed a new manufacturing concept. With no pre-stretching needed anymore, Glare laminates could be produced directly into the shape of the final part, including any thickness steps if so desired. This "composites-like" manufacturing approach enabled large cost reductions compared to the previous "metal-like" approach, thus making Glare parts cost competitive. In addition, the new manufacturing approach allowed integrating local reinforcing elements such as crack stopper bands for very low cost and weight, further improving part performance. Soon after, a third evolution took place, this time resulting in the 7475-based "High Static Strength" (HSS) Glare by further tailoring the material to the specific application requirements. Considering the excellent performance of the Glare parts in the A380 full scale static and fatigue tests, these second- and third generation FML's have indeed successfully bridged the gap from theory to commercial viable operational performance. However, in a dynamic industry like the aerospace industry standing still will soon amount to being taken over by advancing competitors such as Friction Stir or laser beam welded aluminium structures, new alloys such as Al-Li and highly automated CFRP structure fabrication. So, continuous developments are ongoing, taking advantage also of the interesting effect that each improvement in the competing metal and composites industries also inherently offers new potential for better FML’s (new metals, automated composites manufacturing etc.). Several comparative evaluations for costs and performance have been carried out, such as in the European DIALFAST project, as well as in Dutch national R&D programmes (SRP) and in in-house industry studies. These studies and the continuous dialogue with the OEM’s give guidance to research programs to further improve performance and -especially- reduce the cost of FML technology. Recent development topics include - Improved material performance through the use of other aluminum alloys (such as Al-Li) and advanced fibres, such as Zylon - Full scale testing of thin-walled fuselage skins for future single aisle aircraft - Various cost reduction topics, including more automated manufacturing (eg. using tape-layers) We expect that in the near future even more stringent requirements may be set to aircraft structural materials such as better fire resistance and reduced smoke toxicity. This could open new perspectives for FML technology.
Pascal Vermeer, Rene Alderliesten, Rinze Benedictus
Abstract: The current generation aircraft in development (A350 and 787) consists of +50wt% composite as this material couples a low density with extremely high strength and stiffness along the fibre direction. The material generally allows in-plane properties to be freely specified by varying fibre type and changing the ply lay-up, so also a quasi-isotropic state can be obtained. However, a concentrated load transmission strikes on inherent violations of the composite’s basic design; structural coupling by conventional mechanical fastening provokes the notch sensitivity, low shear and bearing strength and the dependence on the laminate configuration. Although increasing the joint thickness would make the design uncritical to bearing, this is in conflict with the aspiration to structural efficiency. In addition, unfavourable effects and increasing complexity are often induced. As fibre metal laminates (FML) proved to be suitable for mechanical joining, adopting the concept of FML seems to be an elegant solution for optimising carbon fibre reinforced polymer (CFRP) joints. The joining area of the CFRP laminate is reinforced by insertion of thin metal foils, replacing those fibre layers whose orientation is contributing least to load transfer. The depth of this reinforced area is only a limited multiple of the edge distance-to-diameter ratio, so that it can be regarded as “local” FML. In most literature, the metallic constituent for a carbon fibre based FML is titanium, hence the designation hybrid titanium composite laminate (HTCL). The transition from the HTCL to the CFRP is the subject of this research, as its design largely affects the cutting sites of the CFRP aircraft structure. Although it seems obvious that the design should be gradual in order to prevent abrupt changes in mechanical behaviour and thermal expansion, its properties and the consequences of different designs on the characteristics of the transition region have not yet been investigated. In order to cover this white spot in research, an experimental programme has been set-up as part of a research project funded by the German Federal Ministry of Economics and Technology with the acronym ENWERUM, which stands for Development of New Materials for Fuselage Applications (translated). The scientific objective of this research is to analyse, characterise and model the mechanical behaviour of the HTCL transition region. The experimental programme involves investigations on: • hybrid titanium composite laminate properties (“full” FML), • (simplified) transition zones under static and dynamic loading, • applicability of specific titanium surface pre-treatments, • response to initial bearing tests. Considering the transition zone, initial experiments suggest the root of the longest titanium ply to be the initiation notch for delamination; the stress concentration due to a bypass mechanism of load accommodates propagation of the delamination front up to a certain level at which the bypass load in a (shorter) neighbouring layer is increased, and a new delamination surface evolves. Enhancement of the experiments by testing with specific equipment (e.g. digital image correlation) and methods (e.g. x-ray computer tomography on delaminated surfaces) on newly produced laminates ought to confirm this theory and the generated data will be input for development of an analytical model for prediction of the persistence of delamination on basis of existing FML/GLARE models. In contrast to classic FML, the HTCL transition will not feature a “through crack”, hence the energy release rate (ERR) approach appears to be the most promising. The study can be considered as an element of advanced hybrid structural concepts in aircraft design and to support the development and certification of the HTCL transition region for actual application by means of adaptation and validation of the ERR approach.
Loris Molent, Simon Barter, Ben Dixon, Ben Main
Abstract: The usage of the F/A-18 aircraft in service with Royal Australian Air Force (RAAF) was revealed to be significantly different to the design spectrum used by the United States Navy for structural certification. The differences lead to further fatigue testing in a joint Canadian and Australian program, using usage representative of both countries in the International Follow-On Structural Test Project (IFOSTP). These tests produced data sufficient to establish the structural integrity basis of both countries’ aircraft. These consisted of the centre fuselage and wing tests (the responsibility of Canada) and an aft fuselage and empennage test performed by the RAAF. Following the centre fuselage test, it was found that the aluminium 7050-T7451 centre barrel (CB) bulkhead’s safe-lives were insufficient (without modification) to meet the required RAAF planned withdrawal date. Thus a CB-replacement (CBR) program was investigated. The Canadians and USN had already commenced a CBR program. For RAAF implementation two main problems were highlighted. The program would be difficult to run in-country and the availability of aircraft during the program would be insufficient to meet the operational needs of the RAAF. For these reasons combined with the predicted expense of such a program the RAAF examined alternative strategies. Additionally, to mitigate against possible uncertainties in the CBR program that were not accounted for in the IFOSTP (e.g. the potential onset of wide spread fatigue damage, additional failure locations, in-service induced defects etc) a teardown and inspection of several ex-service CB’s was conducted by DSTO. To increase the probability of detecting existing “sub-critical” cracks accelerated fatigue testing of the CBs was conducted. This involved the application of representative cyclic loads to the retired CBs in a test rig. Loading was of sufficient magnitude and duration to ensure that potentially life-limiting cracks would be grown to a size that would ensure their detection. This program also allowed a re-examination of the earlier assumptions used to life the CB (e.g. critical crack lengths etc). This resulted in a reduction of the number of CBRs required and thus significant savings of cost and increased aircraft availability for the RAAF. To date, 12 CBs have been tested and torn down including two from RAAF aircraft Teardown inspection consisted of dismantling the CBs to the part level, followed by a detailed inspection of all areas of structural significance. In many cases, quantitative fractography (QF) was performed on the main cracks detected. Data from this program included the type and size of fatigue initiating defects, their locations and distribution, and fatigue crack curves that included in-service growth. It was found that the locations of final failures were very consistent between the CBs tested. This allowed an investigation of the fatigue scatter at discrete locations to be evaluated. This paper briefly summaries the results of the accelerated fatigue testing program, including how these where used to reassess the life of the CBs and evaluate the scatter in fatigue for discrete locations.
Guillaume Delgrange, Rene Alderliesten, Rinze Benedictus
Abstract: For aircraft components made from typical aerospace aluminium alloys and fastened together by rivets or bolts, there are two natural limits which prevent further improvements in fatigue life and damage tolerance: firstly the length of time (the ‘fatigue life’) that aluminium can remain intact under load before cracks appear at stress concentrations caused by fastener holes, and secondly, the relatively fixed ability of the microstructure of aluminium alloys to retard the speed of propagating cracks (a measure of ‘damage tolerance’). There are numerous ways of extending the fatigue life and damage tolerance of such structure; two of the more promising solutions involve firstly adhesively bonding aluminium parts together in order to remove holes and improve fatigue life, and secondly through bonding additional layers of other materials, such as glass-fibres, to aluminium (thus creating a ‘hybrid’ material) to reduce the crack opening effects normally experienced in monolithic aluminium alloys under load. The object of the research is to better understand the fatigue and damage tolerance behaviour of bonded metallic structure and hybrid metallic materials in order to establish the criteria by which these solutions may be designed in preference to conventional aluminium alloy structures. The typical failure associated with these solutions is known as delamination. So far, methods to predict simple in-plane crack opening (Mode II failure) for one-dimensional delamination under shear in thin bonded and hybrid materials exists. However as material thicknesses increase either by additional bonding of material layers or by bonded joints between components, a tensile mode of failure (Mode I failure) is naturally induced. The understanding of the contribution of both these modes of failure to delamination will greatly increase the ability of a designer to specify materials and structure that will prevent delamination from occurring. This paper describes the investigations that have been done to understand the contribution of Mode I and II failure in bonded metallic structure and hybrid metallic materials. Double cantilever beam opening test (pure mode I) and 3 points bending test (pure mode 2) have been performed on specimens with aluminium & glass fibre/epoxy interfaces. Both static and fatigue loading have been tested. For the fatigue loading, the delamination growth rate has been recorded. The data extracted from these tests will be confronted to analytic models such as the one which links the strain energy release rate to the delamination growth rate by a Paris-type equation. The purpose is to get a model that allows the prediction of the delamination behaviour for a specific geometry and loading. Finally, the potential of the developed model, for example the application to finite element models and the extension from basic specimen to a more global structure, will be discussed.
Jim Harrison, Lloyd Hackel, Mike Leap, Jon Rankin, Serena Marley, Joe Nemeth
Abstract: Using laser peening at the highly stressed location where the arresting shoe attaches to the hook shank we have demonstrated a fatigue life enhancement of more than 250% in large scale, geometrically similar fatigue coupons made from Hy-Tuf steel. During an arrested landing on a carrier, the heavy load transferred through the cable to the aircraft induces a combination of axial and bending stresses in the fillet radii near the end of the hook shank. In this application, we employ laser peening to induce compressive residual stress to depths of 0.100 inches (2.5 mm) to the fatigue prone notched end of the shank. When this intense and deep compressive stress is superimposed with the heavy loading stress it reduces the highest surface and near surface tensile loading and significantly improves the fatigue lifetime of the laser peened components. Conventional shot peening, which is standard practice for the production hook shank, was also evaluated in this study.