Browsing by Author "Irving, Phil E."
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Item Open Access AFM observation of surface topography of fibre Bragg gratings fabricated in germanium-boron codoped fibres and hydrogen-loaded fibres.(Elsevier Science B.V., Amsterdam., 2002-11-01T00:00:00Z) Wei, C. Y.; Ye, Chen-Chun; James, Stephen W.; Irving, Phil E.; Tatam, Ralph P.This paper reports the measurement of the surface topology of optical fibres containing a fibre Bragg grating (FBG) using an atomic force microscope (AFM). The AFM observation was made on FBGs fabricated via the phase mask technique in germanium–boron codoped optical fibres, in hydrogen-loaded germanium–boron codoped fibres and in standard telecommunications optical fibres. The surface images reveal that a spatial corrugation pattern was induced by the UV- irradiation, with a period that is half of the period of the phase mask. This UV-induced surface structure was found only on the side of the fibre facing towards the incident UV-irradiation and did not appear on the rear surface. The AFM probe scanned a 10×10 μm2 surface area at seven sites along the 6.0 mm length of fibre that was exposed to the UV-irradiation. The amplitude of the spatial corrugation pattern observed on the AFM image was quantified for each site. It was found that the amplitude in a range of 0.7–3.2 nm was a function of UV-laser intensity distribution and the type of fibre. Hydrogen loaded optical fibres exhibited a corrugation with an amplitude twice as large as that observed in the Ge–B doped fibres that were not hydrogen-loaded. This correlates with the increase in photosensitivity produced by the hydrogen loading. A similar UV- induced spatial corrugation was also observed on standard telecom fibres, but without inducing the refractive index change in the fibre core. The observation of surface topology provides an insight into the structural changes induced during FBG fabrication. UV-induced densification and laser ablation could account for the formation of the surfacItem Open Access Aircraft crash survivability from viscous injury in vertical impacts(Cranfield University, 2009-04) Barth, Thomas H.; Irving, Phil E.This research investigated viscous injury from vertical impact loading to determine if it is critical to survivability of aircraft accidents. A unique database was built from autopsy reports and accident investigations combining injury data with the vehicle impact data. Computer models were created and used to assess injury potential. Common design limits and actual crash data from full scale research experiments were used as inputs. The results were analyzed according to published injury thresholds and compared with real accident autopsies to determine the validity of the hypothesis. Heart and Aortic Injury (HAI) has been considered a critical survivability factor through out the history of mechanized transportation. The mechanisms of HAI in the aircraft environment were never well characterized. Automotive research identified important HAI injury mechanisms related to the forward and lateral impact vectors. This research investigated the vertical impact vector. A model was developed to evaluate the biomechanical response of a simplified visco-elastic system, and incorporated into a system model which included the occupant and aircraft seat. This approach was similar to the development of spine injury criteria and provided the advantage of a macro level evaluation of the injury thresholds and assessment of the criticality in survivable accidents. Evaluations of real accidents sustaining HAI characterized a range of impact severity and approximate boundaries for survivability with HAI and internal organ injury. Viscous injury potential from vertical impact was found to be less critical than potential spine injury. Detailed analysis of HAI documented in autopsy reports and the corresponding accident investigations found that HAI was associated with cockpit environmental factors rather than inertial displacement mechanisms. Vertical displacement of the heart due to inertial loads is not a critical factor in survivable accidents given current aircraft technology. Inertial loading to the heart and aorta is a contributory factor for viscous injuries in aircraft accidents.Item Open Access A comparative experimental study on the diagnostic and prognostic capabilities of acoustics emission, vibration and spectrometric oil analysis for spur gears.(Elsevier Science B.V., Amsterdam, 2007-01-01T00:00:00Z) Tan, Chee Keong; Irving, Phil E.; Mba, DavidPrognosis of gear life using the acoustic emission (AE) technique is relatively new in condition monitoring of rotating machinery. This paper describes an experimental investigation on spur gears in which natural pitting was allowed to occur. Throughout the test period, AE, vibration and spectrometric oil samples were monitored continuously in order to correlate and compare these techniques to natural life degradation of the gears. It was observed that based on the analysis of root mean square (rms) levels only the AE technique was more sensitive in detecting and monitoring pitting than either the vibration or spectrometric oil analysis (SOA) techniques. It is concluded that as AE exhibited a direct relationship with pitting progression, it offers the opportunity for prognosis.Item Open Access Damage sensing in CFRP composites using electrical potential techniques(Cranfield University, 2004-02) Angelidis, Nikolaos; Irving, Phil E.This Thesis investigates the damage sensing capabilities of the electrical potential measurement technique in carbon fibre reinforced polymer composites. Impact damage was introduced in multidirectional laminates and its effect on potential distribution studied. It was found that delaminations and fibre breakages within the laminate can be detected and located by measuring potential changes on the external composite surface. The extent and size of potential changes were significantly affected by the position of the current electrodes in relation to the potential measurement probes. A numerical model was developed investigating the effect of different size delaminations, located in various positions within the lamina, on electrical potential distributions on the external ply, and a quantitative analysis of the numerical results is presented. The numerical simulations demonstrated that the measured potential changes on the external ply were in proportion to the delamination size. The numerical and experimental results were compared and the optimum configuration of current electrodes and potential probes for damage detection selected. The response of electrical potential to mechanical strain, in unidirectional and multidirectional samples was also investigated. It was found that the conductive medium, used for introducing the current, defines the piezo-resistance performance of the composite. A finite element model was developed able to predict the effect of inhomogeneous current introduction in unidirectional specimens on electrical potential and piezo-resistance. The effects of temperature and water absorption on potential measurements were also presented.Item Open Access Damage tolerance characteristics of carbon fibre composites: behaviour under impact loads and post-impact fatigue.(2018-01) Xu, Fan; Liu, Wenli; Irving, Phil E.One of the critical issue restraining the fully application of carbon fibre reinforced polymer (CFRP) composites to aircraft structures is barely visible impact damage (BVID) caused by low-velocity impact (LVI). Consequent internal damage, such as delamination, which is difficult to find by regular inspection, raising a great concern for the damage tolerance performance of CFRP. The main objective of this research study was to investigate the damage tolerance behaviour of carbon fibre composites under impact loads and post-impact fatigue. The project aims to explore and identify delamination growth and failure processes in impact-damaged carbon fibre epoxy composites under compressive cyclic loading and to explain the behaviour in the fracture mechanics method. Low velocity drop weight impacts were used to create the initial damage. There were four levels from 12J-25J. Maximum compressive loads during fatigue were between 70% and 80% of the nominal residual strength in compression after impact (CAI) tests. Delamination propagation was monitored at intervals during the test using DIC and C-scan techniques. It was found that compressive fatigue failure can be categorised into three phases: (1) local bending at impact damage site (2) local buckling mode change and (3) buckling propagation. Monitoring of local normal displacement at the impact damage site provided earlier indications of fatigue induced changes in delamination buckling than observations of delamination area growth. A three-dimension analytical model was developed and revealed that deep delamination has higher strain energy release rates at crack tip than delaminate at the surface of laminate.Item Open Access Damage tolerance of welded aluminium aircraft structures(Cranfield University, 2000-11) Bussu, G.; Irving, Phil E.Riveting is a traditional joining technique mostly used in the manufacturing of aircraft structures. Manufacturing studies on the next generation of wide body commercial aircraft have' indicated that the achievement of acceptable cost/benefit goals would require the» application of highly cost-effective joining processes. Although riveting provides good structural performance, it is expensive and time consuming. Welding is a candidate process t be used to manufacture large aircraft structures allowing sensible cost reductions and structural efficiency. Welded aluminium 2024-T351 structural joints produced .with a new generation of welding processes, such a Friction Stir Welding ( SW) and Plasma welding were characterised in terms of rnicrostructure, hardness and weld residual stress. Tensile properties and stress-strain behaviour of the FSW joints was investigated and discussed using simple mechanical models. The investigation of the fatigue properties of the FSW and Plasma Welded structural joints revealed the superior behaviour of the FSW joints. It was found that fatigue strength in FSW joints is dominated by surface irregularities produced by the welding process. Weld surface skimming greatly improved fatigue strength by removing surface stress concentrations. Initiation in skimmed joints occurred at locations of minimum hardness. Fatigue endurance behaviour of skimmed joints was equal or superior to that reported in riveted aluminium joints. Fatigue crack propagation studies were carried out on FSW 2024-T351 joints for cracks parallel and orthogonal to the weld direction. Crack propagation behaviour was sensitive t both weld orientation and the distance of the crack from the weld line. Growth rates both faster and slower than in the parent material were observed, depending on the crack orientation and distance from the weld. Residual stress was mechanically relieved and the effects on crack propagation observed. A comparative analysis of the results associated with crack closure measurements indicated that crack growth behaviour in the FSW joints was generally dominated by the weld residual stress. Possible design solutions for damage tolerant Welded aircraft structural components were identified and discussed in the light of the experimental results.Item Open Access Development of fatigue cracks from mechanically machined scratches on 2024-T351 aluminium alloy - Part II: finite element analysis and prediction method(Wiley, 2016-11-10) Cini, Andrea; Irving, Phil E.A prediction method to evaluate the effect of scratch geometry on fatigue life of aluminium structures containing scribe marks was developed on the basis of the experimental results described in Part I of this paper. Finite element calculations were performed on scribed samples to investigate the local stress around scribes. Elastic and elastic plastic stress and strain distributions at the scribe root were computed under monotonic and cyclic tensile and bending loads evaluating the driving force behind initiation and propagation from scribes. Scribe shape, size and cladding regulated stress and strain distributions in the neighbourhood of scribe roots. Fatigue life of tested scribed samples was divided into initiation life, defined as the cycles spent to develop a 50 μm deep crack at scribe roots, and the remaining propagation life up to failure. Striation counting measurements were used to calculate propagation lives by integrating linear elastic da/dN vs. ΔK curves. Only up to a maximum of 38% of total fatigue life was spent to propagate an initial 50 μm deep crack from scribe roots. The theory of critical distances was successfully applied to predict initiation lives of scribed samples from elastic stress distributions. A plastic correction was also suggested in the frame of the theory of critical distances, to correlate initiation lives of clad and unclad specimensItem Open Access Development of fatigue cracks from mechanically machined scratches on 2024-T351 aluminum alloy - Part 1: experimentation and fractographic analysis(Wiley, 2016-10-27) Cini, Andrea; Irving, Phil E.Clad and unclad 2024-T351 aluminium alloy sheets, weakened by mechanically machined scratches, were fatigued to investigate the effect of small surface damage, like scribe marks, on aircraft fuselage joints. The role of scratch cross section geometry on fatigue life of scribed components was analysed. Scratches between 25 and 185 µm deep, with 5, 25 and 50 µm root radii, were cut on sample surface by using diamond-tipped tools. After testing, fracture surfaces were examined using a scanning electron microscope, and crack growth rates were measured by striation counting. Scratches reduced aluminium fatigue life under tensile and bending load up to 97.8% due to multiple crack nucleation at their roots. Short cracks nucleated from sharp scratches coalesced to form unique elongated cracks growing through sample thickness. Cracks initiated from scratches were typical short cracks, growing faster than conventional long cracks. Despite the different scribing process, fatigue data of regular diamond tool cut scribes can be used to conservatively predict life reduction owing to ploughed in-service scribe marks on fuselage joints. Finite element analyses on scribed samples and the fatigue life prediction models are described in Part II of this paper.Item Open Access Development of probability of detection data for structural health monitoring damage detection techniques based on acoustic emission(Stanford University, 2013-12-12) Gagar, Daniel; Irving, Phil E.; Jennions, Ian K.; Foote, Peter; Read, Ian; McFeat, JimStructural Health Monitoring (SHM) techniques have been developed as a cost effective alternative to currently adopted Non-Destructive Testing (NDT) methods which have well understood levels of performance. Quantitative performance assessment, as used in NDT, needs to be applied to SHM techniques to establish their performance levels as a basis for technique comparison and also as a requirement for practical aerospace application according to set regulations. One such measurand is Probability of Detection (POD). This paper reports experiments conducted to investigate the location accuracy of the Acoustic Emission (AE) system in monitoring events from HsuNielson and fatigue crack AE sources as a route to establish the POD of AE in SHM. It was found that fatigue crack tips could be located at 90% POD within 10 mm accuracy.Item Open Access Diagnostics and prognostics with acoustic emission, vibration and spectrometric oil analysis for spur gears; a comparative study(Learned and Professional Society Publishers, 2005-08-01T00:00:00Z) Tan, Chee Keong; Irving, Phil E.; Mba, DavidWhilst vibration and spectrometric oil analysis for gear fault diagnosis are well established, the application of AE to this field is still in its infancy. This paper describes an experimental investigation on spur gears in which natural pitting was allowed to occur. Throughout the test period, AE, vibration and spectrometric oil samples were monitored continuously in order to correlate and compare these techniques to natural life degradation of the gears. It was observed that the AE technique was the most sensitive in detecting and monitoring pitting.Item Open Access The effect of microstructural parameters on the mechanical properties of non-crimp fabric composites(Cranfield University, 1996-08) Miller, A. J.; Irving, Phil E.Textile composites are being considered for use in aerospace primary structures as they have the potential to reduce the cost of manufacturing composite structures. Non-Crimp Fabrics (NCFS) are one class of textile composites which are being investigated for this use. The aim of this project was to determine what affect the fabric structure of the NCF has on the larninates microstructure and in turn what affect this has on the final mechanical properties. After an explanation of the need for textile composites a review of the advantages and disadvantages of the different classes is presented. The concept of a link between the microstructure of the composite and its mechanical properties is introduced with a review of past work in this area. Image analysis methods are identified as the most promising microstructural measurement techniques and their past use in quantifying composite microstructures is surveyed. Methods of quantifying the tow crimp, resin rich areas and fibre orientations were developed using image analysis techniques. These methods were then used to quantify the microstructure of a wide range of NCF laminates, with woven and unidirectional materials being studied for reference. It was found that the initial structure of the NCF had an influence on the final nicrostructure of the laminate. Both the amount of tow crimp and resin rich areas were affected by the spacing of the tows, stacking sequence, stitch density and stitching material. It was also shown that the fibres inside the tows had large orientation variations across their cross-section. The mechanical properties of the materials wer'e evaluated by compression and Interlaminar Shear (ILS) testing. It was found that that an increase in the tow crimp reduced the compression strength while an increase in the resin layer thickness decreased the IS strength. It was shown that these results agreed with past experimental and modelling work.Item Open Access The effect of sequential action of corrosion and fatigue on fatigue crack initiation and propagation in 2024-T3 aluminium alloy(Cranfield University, 1997-09) Roungos, I. C.; Irving, Phil E.The influence of sequential action of corrosion and fatigue on fatigue behaviour of 2024- T3 aluminium alloy was investigated. EXCO solution was used for corrosion development. Corrosion evolution and penetration were investigated in terms of exposure time, dimensional characteristics, orientation, microstructure of material, incremental corrosion application and sequential application of corrosion and fatigue. Single Edge Notch Tensile specimens were used in the fatigue-corrosion tests. Fatigue intervals were interrupted by corrosion application. Four fatigue intervals and four corrosion segment times were incorporated in the test matrix. The results of fatigue ínitiation, propagation and total life were compared to the virgin and pre-corroded materials' behaviour. The main mechanisms of crack acceleration and arrest were identified and discussed relative to the morphology of corrosion development. Finally a comparison was tried between the experimental lives and the predicted ones, calculated from two crack growth packages using the most conservative approach.Item Open Access Effect on Fatigue Performance of Residual Stress induced via Laser Shock Peening in Mechanically Damaged 2024‐T351 Aluminium Sheet(Cranfield University, 2014-12) Smyth, Niall; Irving, Phil E.During manufacture and maintenance the fuselage skin of aircraft are susceptible to damage in the form of scratches. Normally not considered to be of major concern to aircraft structural integrity some airlines discovered fatigue cracks had initiated at the root of scratches. Crack propagation was in the through thickness direction and if left untreated could cause rapid decompression of the passenger cabin. Standard repair methodology requires patches be riveted around scratch damage and in extreme cases could require entire replacement of affected skin panels. Laser shock peening (LSP) is an emerging surface treatment that has been shown to improve fatigue performance of safety critical components by inducing a surface layer of compressive residual stress. In this work LSP was applied along the scratch damage in an effort to restore pristine fatigue performance. The aim of the project was to model the effect on fatigue crack growth rate of residual stress fields induced via LSP and to validate predictions by comparison to experimental test results. The scratches were recreated under controlled laboratory conditions using a diamond tipped tool. This process allowed creation of reproducible V shaped scribes to controlled depth, wall angle and root radius. Scribes of depth 50 and 150 μm with root radius 5 μm were created in dogbone shaped samples of 2 mm thick 2024‐T351 clad aluminium. Samples were tested in fatigue at an R = 0.1 and maximum stress of 200 MPa. The scribe damage reduced fatigue life compared to the pristine material by a factor of 22. Scribed samples were processed using LSP treatment from different providers that created known residual stress fields in the material. The fatigue life of scribed samples after peening varied from a further decrease to a 13 times increase dependent on the residual stress field induced. An elastic‐plastic crack closure based finite element model was created to determine the effect on stress intensity factor and stress ratio of residual stress. Fatigue lives calculated were within a factor of 2 of experimental lives. It was predicted that crack closure was present during up to 80% of the applied load cycle due to the compressive residual stress field. However plasticity induced crack closure actually reduced after peening because the compressive residual stress field induced a smaller plastic zone at the crack tip and hence reduced the plastic wake. A residual stress based fatigue life sensitivity study was performed to optimise the profile of the residual stress field for improved fatigue performance. The required profile was created in test samples using LSP. The fatigue life of peened samples increased by a factor of up to 15 however pristine life was not fully recovered. A restriction imposed by the industrial application was peening applied to one face only. This created an unbalanced stress field that resulted in sample distortion to maintain equilibrium. The distortion induced out of plane bending stresses during testing and caused premature crack initiation on the unpeened face. However using interrupted fatigue tests it was found that although crack initiation also occurred at the root of the scribes the cracks were arrested after 24 μm of propagation. This was consistent with the findings of the crack growth prediction model.Item Open Access The Effects of Residual Stress and HAZ on Fatigue Crack Growth in MIG Welded 2024 and 7150 Aluminium(2003-01-01T00:00:00Z) Lin, J.; Ganguly, Supriyo; Edwards, L.; Irving, Phil E.The effect of weld residual stress, and hardness profile on fatigue crack propagation in MIG welded 2024 and 7150 aluminium joints was studied. Residual stress fields in the weld were measured using neutron diffraction. Tests were performed using a range of mean stresses and on different starting defect shapes. It was found that fatigue crack propagation in weld metal and HAZ is influenced by crack morphology and also the distribution of residual stress and hardness. Control of these parameters will optimise damage tolerance capability in welded safety critical structures.Item Open Access An experimental approach to quantify strain transfer efficiency of fibre bragg grating sensors to host structures(2001-06-25T00:00:00Z) Wei, C. Y.; Ye, Chen-Chun; James, Stephen W.; Tatam, Ralph P.; Irving, Phil E.This paper developed a method to evaluate the strain transfer efficiency of fibre Bragg grating sensors to host structures. Various coatings were applied to fibre Bragg grating sensors after being fabricated. They were epoxy, silane agent and polypropylene, representing different surface properties. A neat epoxy resin plate was used as the host in which the coated fibre sensors were embedded in the central layer. The tensile strain output from the FBGs was compared with that obtained from electrical strain gauges which were attached on the surface of the specimen. A calculating method based on the measured strains was developed to quantify the strain transfer function of different surface coatings. The strain transfer coefficient obtained from the proposed method provided a direct indicator to evaluate the strain transfer efficiency of different coatings used on the FBG sensors, under either short or long-term loading. The results demonstrated that the fibre sensor without any coating possessed the best strain transfer, whereas, the worst strain transfer was created by polypropylene coating. Coatings play a most influential role in strain measurements using FBG sensors.Item Open Access Extending fatigue life of aircraft fuselage structures using laser-peening(2017-10) Busse, David Osman; Irving, Phil E.; Ganguly, SupriyoFatigue of airframe structures is a constant challenge to aircraft manufacturers when designing, maintaining and repairing new and aging metallic components. Laser-Peening (LP) is a highly flexible and controllable surface treatment and relatively new to manufacturers of large civil aircraft which demonstrated that it can extend the fatigue and crack growth life in aluminium alloys by introducing deep compressive Residual Stresses (RS). Currently there is no application of LP to any components of large civil aircraft. The aim of this research was to demonstrate and explore different LP strategies that can produce significant extension of the fatigue and crack growth performance of aircraft fuselage structures using Laser-Peening. Two representative samples made from 2000 series aluminium alloy were designed to represent features of the fuselage: A Centre Cracked Tension (CCT) panel made of 1.6 mm thick 2524-T3 represented the fuselage skin. Single overlap Lap-Joints (LJ) of 2.5 mm thick 2024-T3 aluminium with titanium Hi-Lok bolts arrayed in 5 columns and 3 rows embodied longitudinal LJ of aircraft fuselages. Both test samples were laser-peened without protective coating (LPwC) using a range of LP strategies in which LP process parameters and spatial arrangements of laser-peened areas were systematically varied. RS fields were measured before fatigue testing under constant amplitude loading. RS measurements used Incremental Centre Hole Drilling (ICHD) and X-ray and Neutron diffraction techniques. Laser-peening produced peak compressive RS of 200 – 350 MPa and compression stress penetration depths between 700-1000 μm. These values are superior to RS profiles induced by Shot-Peening. The value of peak compression stress and penetration depth depends on LP process parameters and on the LP layout. The latter defines the location and size of the laser-peened areas. A study of the effect of different LP strategies to establish the most effective LP treatment to enhance crack growth life of fuselage skins was performed using a Finite Element based crack growth model. The model was first used to introduce balanced RS fields into a cracked CCT sample. The effective stress intensity factor range (ΔKeff) and effective R-ratios (Reff) were then calculated as the crack tip progressed through the sample. Subsequently, fatigue crack growth rates and lives were computed using Walker’s empirical crack growth law. The accuracy of the model was demonstrated by comparison with crack growth test results from laser-peened CCT-samples. Results of the parameter study showed that an increase in the level of compression within the LPS increased life most significantly. Increased width of peen stripe increased the life while increasing the distance of the stripe from the starting position of the crack tip decreased the life. Four different LP strategies were applied to LJ samples. Subsequent fatigue testing demonstrated fatigue life improvements of between 1.14 to 3.54, depending on the LP strategy. The LP layout was identified as a key parameter determining the fatigue life. It was found that when small LP areas were used, to leave as much elastic material as possible between the peened areas, larger compressive stresses and minimised balancing tensile stresses were produced. Observations of fatigue fractures on joint samples showed that crack initiation occurred remote from the fastener holes, either in regions of fretting fatigue in peened areas or in regions of balancing tensile stress adjacent to peen boundaries. Optimum fatigue lives occurred when both fracture types occurred in the same sample. Striation spacing measurement and analysis showed that compressive residual stresses had little or no effect on fatigue growth rates at crack lengths < 600 µm. The majority of fatigue life extension was achieved during initiation and crack growth < 600 μm. The obtained results established evidence of how aircraft fuselage structures made of conventional 2000 series aluminium-copper alloys can be effectively laser-peened to produced fatigue life improvements and also of how to avoid any detrimental reductions in fatigue life which can also occur when LP is applied randomly. The generated research conclusions are applicable to other metals, geometries and components.Item Open Access Fail-Safe Design of Integral Metallic Aircraft Structures Reinforced by Bonded Crack Retarders(Elsevier Science B.V., Amsterdam., 2009-01-01T00:00:00Z) Zhang, Xiang; Boscolo, M.; Figueroa-Gordon, Douglas J.; Allegri, Giuliano; Irving, Phil E.This paper presents an investigation on the effectiveness of crack growth retarders bonded to integral metallic structures. The study was performed by both numerical modelling and experimental tests. It focuses on aluminium alloy panels reinforced by bonded straps made of carbon-epoxy, glass-epoxy composite materials or a titanium alloy. The goal was to develop a fail-safe design for integrally stiffened skin-stringer panels applicable to aircraft wing structures. The modelling strategy and finite element models are presented and discussed. The requirements that the models should meet are also discussed. The study has focused on establishing the extent of crack retarder benefits, in terms of fatigue crack growth life improvement, by numerical simulation and experimental tests of various crack retarders. The results of predicted fatigue crack growth retardation have been validated by tests of laboratory samples. This study concludes that by bonding discrete straps to an integral structure, the fatigue crack growth life can be significantly improved.Item Open Access Fatigue crack growth rates under variable amplitude load spectra containing tensile underloads(Cranfield University, 2003-10) Zitounis, Vasilios; Irving, Phil E.An extensive research program was performed to investigate the load interaction effect of the combined action of small amplitude high R ratio cycles and large amplitude low R ratio underloads on the crack growth of large cracks. The study was driven by the needs of the damage tolerance approach in the helicopter structures, which requires robust knowledge on the crack growth behaviour of the advance high strength alloys under the characteristic helicopter spectra loading. The study was conducted on three metallic alloys, Ti-10V-2Fe-3Al, Al8090 T852 and Al7010 T76351 using compact tension specimens (w=70mm, t=17mm). The potential drop technique was used for the measurements of the crack length. The crack opening loads were determined from the applied load versus crack opening mouth displacement curve using a curve fitting technique and crack opening displacement gauge. The experimental results show that cracks can grow faster than the life predictions with no load interaction effects under spectra containing tensile underloads. The acceleration effects are different depending on the number of the small cycles, the Kmax, the R ratio of the small cycles, the underload cycle and the material. Significant closure observations on the underloads and on the small cycles of variable amplitude loading spectra were made. Based on the test finding and on the studies of other researchers, it is suggested that the acceleration effects are mainly due to the reduction of crack opening point of the tensile underloads comparing with the Constant Amplitude Loading (CAL) data. An extensive evaluation of the ability of FASTRAN model to predict the fatigue lives under the tested loading spectra was carried out. The evaluation focuses on the influence of the constraint factor a and the ∆Keff curve inputs on the predictions. The model produces very good and consistent predictions for the three alloys, when the inputs represent adequately the actual fatigue mechanism. The model predicts the measured acceleration effects by reducing the closure level of the underloads.Item Open Access Fatigue cracking behaviour of epoxy-based marine coatings on steel substrate under cyclic tension(Elsevier, 2017-02-13) Wu, Tongyu; Irving, Phil E.; Ayre, David; Jackson, P.; Zhao, F.Strain controlled fatigue tests have been performed on two types of heavily filled epoxy corrosion protection coating sprayed onto a 6 mm steel substrate. Fatigue cycling was performed at R ratios of 0 and −1. The two coatings differed in their formulation and the major differences in mechanical performance were in their static strain to first crack development and their fracture toughness, where Coating A was significantly tougher than coating B. During strain cycling coating crack development was monitored using optical observations and surface replicas. It was found that in both coatings surface crack development began soon after the onset of cycling and proceeded via growth of surface channelling cracks and multiple initiation of new cracks. Detailed studies were made of crack development morphology and its relation to coating type and to the applied strain range. A definition of coating life as the first appearance of a 2 mm surface crack length was used. This represented the end of the life where the coating protected the substrate. Before this life was achieved, crack growth rates of single cracks were invariant with crack length. After this point further crack growth, multiple cracking and crack to crack interactions took place. Cracking in this region could be characterised with a new total crack length parameter shown to be strongly dependent on applied strain range.Item Open Access Fatigue life enhancement of aircraft structures through bonded crack retarders (BCR)(Cranfield University, 2015-09) Doucet, Jeremy; Irving, Phil E.; Zhang, X.The trend in aircraft design is to produce greener airplanes through lighter structures and/or structures with extended life and reduced maintenance. Bonded crack retarders (BCR) are one of the solutions towards that objective. BCR are reinforcing straps bonded to the structure in order to improve the fatigue and damage tolerance properties of the assembly. The aim of this study was to demonstrate that the BCR hybrid technology – beneficial for upper wing cover – could also be applied to lower wing covers. The project also focused on evaluating BCR most important parameters. The fatigue life improvement obtained from BCR was evaluated through a series of coupons and skin-stringer assemblies tested under constant and variable amplitude loading. While the coupon tests demonstrated a life improvement of only 17% under constant amplitude loading, the variable amplitude load tests performed on the skin-stringer assembly demonstrated increased fatigue lives with a factor of 5 and reduced crack growth rates with a factor of 5 to 6. A finite element calculation tool was developed in order to conduct a parametric analysis of BCR geometry through the evaluation of the substrate stress intensity factor in the case of fatigue loading. The main difficulty was to include the interacting mechanism of the substrate lead crack and the disbond of the adhesive layer. The novelty of the approach was to incorporate the fatigue delamination calculation in order to evaluate the fatigue disbond propagation with crack growth. This was embedded in a 3D finite element design tool ReSLIC (Reinforced Structures Life Improvement Calculation). A necessary step to the development of ReSLIC was the analysis of fatigue properties of the adhesive system in order to provide input data for fatigue delamination calculations. To that end, a series of fatigue tests were performed in pure Mode I, pure Mode II and mixed mode with ratios of 25%, 50% and 75% of mode II ... [cont.].
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