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Poster pitches day 2
Chair: Marcel Bos
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Towards understanding residual strength and damage evolution in damaged composite laminates
John-Alan Pascoe, Wenjie Tu, Davide Biagini, René Alderliesten
Abstract: Composite materials are increasingly used in aerostructures, making up about 50% of the structural weight of the latest airliners, and up to 90% in helicopters. Managing the damage such structures will inevitably sustain is therefore increasingly important in both design and sustainment. Understanding the significance of damage findings requires answering two questions: (1) what is the impact on the residual strength, and (2) how will the damage evolve under fatigue loading?
Unfortunately, prediction tools to answer these questions remain lacking. Available models mainly rely on experimental data gathered during development of the aircraft, e.g. to substantiate that the design meets the ‘no-growth’ criterion. However, the complexity of the damage mechanisms and their lay-up dependence, makes it difficult to generalise these data, to predict the behaviour of damage scenarios that don’t precisely match the test conditions. This results in the application of very strict, and likely over-conservative, damage limits to composite structures.
We will show that one of the reasons for current limitations of predictive models, is that damage characterisation is driven more by the limitations of inspection techniques, than by an understanding of the physical damage processes. Additionally, investigations into fatigue delamination growth rely on specimens that are not representative of actual composite structures. We will present ongoing efforts in our lab to address these challenges. This includes using acoustic emission sensors to distinguish damage mechanisms during quasi-static and fatigue compression after impact testing, measuring of fatigue delamination growth in multidirectional interfaces, and using pulse-echo ultrasonic scanning to build up a 3D image of the delamination configuration and propagation, under either in-plane or out-of-plane loading.
Based on the results we conclude:
• Projected delamination area or delamination width are insufficient to properly quantify damage severity; a more detailed damage description is necessary.
• Planar (2D) delamination growth in multidirectional laminates involves delamination migration processes that are similar to those seen in 1D delamination tests in multidirectional laminates, but which are not represented in standard unidirectional delamination growth coupons.
• Further research is necessary to understand what data is needed to enable accurate prediction of residual strength and remaining life when damage is detected.
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Experimental study on the effect of different fatigue loading method on the property of metal cantilever beam
Wei Cui, Dingni He, Jianjian Yu, Yunfei Liao
Abstract: In order to investigate the influence of different fatigue loading methods on the property of metal cantilever, the resonant fatigue testing system and the corresponding conventional fatigue testing system were designed respectively based on the electromagnetic vibration table, and the bending fatigue tests were conducted on LY12CZ aluminum alloy cantilever beam specimens. A special progressive analysis based on damage equivalence was used to obtain S-N curves under resonance environment in the range of 1×105-1×107 cycles. The results show that the fatigue limit of LY12CZ aluminum alloy under bending resonance environment is 74.6MPa and the fatigue limit under conventional condition is 76.2MPa at 1×107cycles. When the number of cycles is less than 5×106, the bending resonant fatigue property of LY12CZ aluminum alloy is significantly different from the conventional fatigue property, while in the range of 5×106-1×107cycles, there is little difference between the two properties.
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Research on structural damage identification methods for aircraft full scale fatigue test
Shaozhen Pan, Guiyong Zhong, Xiaodong Liu
Abstract: In aircraft full scale fatigue test, identify and locate structural fatigue cracks accurately is an important job, which traditionally was completed by regular Non Destructive Inspection (NDI). While, due to the high complexity of structure geometry and the limitation of detection space, fatigue cracks found by NDI tend to be larger in size or even missed in detection, which is extremely disadvantageous for structural repair.
In order to solve the above problems, several methods such as threshold method, relative error method, probability distribution method and slope method have been presented. However, it is found that these methods have insufficient ability to identify short cracks and poor adaptability to strain data of different masses in the application of several full scale fatigue tests.
In view of the above problems, a strain sensor based method on fatigue damage monitoring and identification for critical locations of aircraft structure in full scale fatigue test was established, on the foundation of high accuracy stress modelling analysis. Firstly, an approach used to distinguish failed strain sensors was presented in order to avoid false damage warning. Secondly, a data cleaning method which combined K-means clustering algorithm with 3σ rule was developed and used to detect and eliminate outliers in sensor data. Lastly, a fatigue damage identification method with adaptive threshold was proposed. In aircraft full scale fatigue test, several fatigue cracks on critical structures were identified and located timely and successfully using these methods. More importantly, it is more accurate and sensitive compared with existing methods, such as threshold method, probability distribution method, relative error method and slope method. Methods presented in this paper were significant for structural design modification and crack repair cost reduction. It can also be applied to structural health monitoring.
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Modal testing of vertical tail of F/A-18 Hornet
Vesa Nieminen
Abstract: The goal of the study was to identify experimentally dynamic vibration characteristics of the aft fuselage and especially the Vertical Tail (VT) of a fighter aircraft. The object of the study was F/A-18 Hornet of the Finnish Air Force. Vibration properties, i.e., natural frequencies, modal damping factors and mode shapes were determined experimentally by impact testing. Special focus was on the modes around about 15 Hz and 45 Hz. Based on earlier analysis of the measured (in-flight) acceleration and strain data, it has been found that the VT experiences high vibration levels around these frequencies due to flow-induced excitations.
The plane was standing on the landing gears during the measurements. The left and right Vertical Tails were excited by an instrumented impact hammer having a soft plastic tip. Both VTs were excited separately and responses from both VTs and other locations were measured during all tests. Frequency Response Functions (FRFs) were calculated between measured input excitation force and acceleration responses. To identify closely spaced double modes, separate measurements were conducted where both VTs were excited simultaneously randomly at random locations of the VT surface by two impact hammers having soft plastic tips. Time histories of acceleration responses were recorded for Operational Modal Analysis (OMA). Natural frequencies, modal damping factors and mode shapes were identified using both conventional experimental modal analysis and Operational Modal Analysis methods.
Lowest elastic natural modes of the VT of the F/A-18 Hornet were identified experimentally successfully. It was found that due to symmetry of the structure, main VT modes are divided into symmetric and antimetric modes having close natural frequencies. Random impact excitation technique for Operational Modal Analysis was demonstrated to be applicable for identification of very closely spaced modes.
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A study on the reliability life evaluation method for aircraft structure details based on PFMA
LI Sanyuan, CHEN Xianmin, DONG Dengke, CHEN Li
Abstract: The aircraft structure includes a variety of typical details which determine the reliability life of aircraft structures. The modern aircraft structure should meet the requirements of long lifetime, high reliability and good economy, etc. Therefore, it is significant to perform the evaluation of the reliability life for aircraft structure details. In this paper, based on the Probabilistic Fracture Mechanics Approach (PFMA), we conduct the following study on the reliability life evaluation method for aircraft structure details.
(1) Normal information diffusion estimation and grey estimation method are introduced to improve the estimation accuracy of the population distribution of the 3-parameter Weibull consistent distribution. The numerical analysis results show that the grey estimation method gives the best accuracy compared with the Gauss-Newton iterative method and the common correlation coefficient method.
(2) In order to definite the calculation formula of crack exceedance, a critical factor of crack growth rate is proposed. Moreover, a reliability life evaluation method is put forward by combining the probabilistic durability with damage tolerance analysis. Furthermore, the Monte Carlo method is introduced to avoid complex integral operations. It is shown that the computing efficiency of Monte Carlo method is much better than the numerical integration method within a certain accuracy.
(3) The reliability life evaluation method is programmed into a software, which makes the complex calculation processes simplified and more convenient for engineers. For the fatigue problem of notch details of the stringer ends in the upper panel of aircraft central wing, the stress level calculation is conducted for the weak details through Finite Element Analysis (FEA). The crack propagation data is obtained by fractographical analysis, as well. It is proved that the reliability life evaluated in this study is in line with that in the full-scale aircraft structural fatigue test.
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Testing of a cocured compsite frame panel by the applications of vacuum as an alternate to frame bending test
Kotresh Gaddikeri, SR Viswamurthy, CH Viswarupachari, BL Dinesh, Rueben Dinakar, Sven Werner
Abstract: The Frame Bending Test (FBT) of fuselage panels is plagued by complex design at load introduction regions, high workload for assembly of specimen to test rig and the need for disassembly for access to stiffened structure. An alternative to the FBT was explored by the application of vacuum on skin side of panel using a metallic fixture while frame side of panel is subjected to atmospheric pressure. The vacuum level can be controlled to obtain the desired differential pressure.
A curved composite panel was designed with three cocured corrugated frames and eight stringers under the InFuSe (Integral Fuselage Shell Concepts) Project between CSIR-NAL and Airbus. The shear clips, to stabilise the frame web laterally, were eliminated by the corrugation of web. A metallic fixture was developed to mount the panel to enable application of vacuum. The finite element (FE) analysis of panel mounted on fixture was carried out to understand the structural response. The desired circumferential strains in panel were achieved by the proper sizing of vacuum fixture. The location of strain gauges, Digital Image Correlation (DIC) regions and dial gauges were guided by FE analysis. Acoustic Emission (AE) was also monitored during the test.
Two vacuum tests were carried out with vacuum levels of 100mbar and 20mbar. The structural responses were measured both during loading and unloading. The first test was stopped at 100mbar vacuum pressure because of increased AE activity. Post-test ultrasonic scan of cocured joints showed disbonds in the Frame and Stringer crossover regions in proximity to metallic fixture. Rivets were installed on disbonds to prevent the further growth. Subsequently, the panel was loaded up to 20mbar vacuum pressure and the panel withstood the vacuum pressure successfully. Ultrasonic scan was carried out on cocured joints showed no disbonds. The structural response in terms of deflections and strains were correlated well with FE simulations.
The proposed Vacuum Test has advantages like smooth and uniform load introduction, quick assembly and economical. It also allows quick access to specimen for DIC, NDE and other sensors on frame side during the test and presents itself as an alternate to FBT.
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Three-dimensional weight function analyses and stress intensity factors for general surface and corner crack emaninating from circular hole
Wu Xu, Bo Zhang, X.R Wu
Abstract: Comprehensive study of aircraft structural failures showed that the most prevalent failure is due to cracks originating from fastener holes, where stress concentration takes place. Stress intensity factors (SIFs) for these crack configurations are the prerequisite for evaluation of the critical crack sizes and fatigue lives of components during which the initial cracks would grow to the critical size. The computational efficiency for SIF is critical, since for every crack growth step, the crack size changes and so are the SIFs. Much work has been done to obtain SIFs of a single and double symmetric 3D cracks at riveted holes subjected to typical loadings. However, due to the complexity of the problem, limited work has been done for the SIFs of the crack configurations as shown in Fig.1 subjected to arbitrary load cases. These crack configurations are very common, but are too complicated to be analytically tackled. The SIF solution is still highly demanded for the damage tolerance analyses of various flaws at fastener hole. In this paper, the 3D slice synthesis weight function method (SSWFM) is further developed to calculate the 3D SIFs for more general and complex case, eccentric and asymmetric surface and corner cracks, and combination of surface and corner cracks subjected to arbitrary load cases. The resulting 3D SIFs are extensively compared to those obtained from FEM/Franc3D; very good agreement is achieved as shown in Fig.2. However, the developed SSWFM is about 450 times faster than FEM/Franc3D in calculation of the 3D SIFs for the present complex crack geometry. It can be used to obtain 3D SIFs of most of the surface and surface-corner crack configurations from a hole in practical engineering, and would be useful for 3D fatigue crack growth analysis of rivet joint structures subjected to various load cases.
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Influence of microstructure on small crack growth behaviour of maraging steel weld
Haitao Zhao, Delon Guo
Abstract: Experiments were carried out to study the influence of microstructure on fatigue small crack growth behavior of 18Ni(250) maraging steel and weld. Microstructure characteristics of base metal and weld metal were analyzed by optical microscope and scaning electron microscope. Results show that the segregation of elements such as Mo, Ti and Ni leads to the formation of massive reversed austenite phases in the weld interdendrites, and correspondingly reduces the content of these elements in martensite matrix of weld metal, thereby reducing the amount of N3(Mo, Ti) in the weld matrix, leading to decrease of microhardness and tensile strength of weld metal compared to that of base metal.. Fatigue tests show that the behaviour of fatigue small crack growth in weld metal is obviously faster than that in base metal. The distribution characteristics of test data of fatigue crack growth rate are similar to that of large crack for base metal. Clearly, there is coalescence of microcracks in small crack growth path of base metal. And this phenomenon rearly occurred in the small crack growth path of weld metal.
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Investigation of the tensile/compressive residual stresses in AISI 4340 steel under low-cycle fatigue loading
Marian Patrick, Jeremy Laliberte, Xin Wang
Abstract: Landing gear fuse pins are notched energy-absorbing components designed to reduce the possibility of aircraft structural damage in the event of a hard landing. The area surrounding a fuse pin notch is in a state of complex stress, causing local yielding under applied loading. This work utilizes the strain-life method to utilize local yielding behaviour to study the low-cycle fatigue life of notched AISI 4340 steel components and to study the effects of complex stresses.
Experimental testing involved mechanical property testing, finite element analysis (FEA) simulation, and low-cycle fatigue testing of notched specimens. Control specimens were subjected to quasi-static low-cycle fatigue using the strain-life model. Variable specimens were subjected to either a quasi-static initial overload prior to fully reversed low-cycle fatigue loading, or a mean stress of identical magnitude as the residual stress induced from the applied overload. The magnitude of the residual stress at the notch root was determined using elastic-plastic FEA in ABAQUS-2019x. It was found that the tensile residual stress induced through compressive overload caused a slight decrease in fatigue life, while the compressive residual stress induced by tensile overload caused a slight increase in fatigue life. The residual stress was approximated as a mean stress to determine the suitability of this approximation as a design assumption from the literature. It was found that applying the tensile residual stress as a positive mean stress causes a significant reduction in fatigue life to provide a conservative life estimate. However, applying the compressive residual stress as a negative mean stress causes a significant increase in fatigue life, or a non-conservative estimate. These results are significant in the design of components where plasticity is not negligible in fatigue life assessments. The significance of the residual stress to mean stress approximation is critical in engineering design to ensure conservative estimates are used in engineering applications
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Influence of the layer thickness on the very high cycle fatigue behaviour of composite materials
Martin Bartelt, Tim Luplow, Peter Horst, Sebastian Heimbs
Abstract: High-efficiency structures like blades of wind turbines and helicopters undergo over 108 load cycles during their lifetime. For composite materials, a lack of research in this so-called very high cycle fatigue (VHCF) regime may lead to conservative designs or fatal failures.
One aspect known for its strong influence on static and fatigue crack initiation is the layer thickness of the composite plies. As the thickness decreases the crack initiation stress grows. The effect of the layer thickness under VHCF promises higher lifetimes and enhanced structures.
To investigate this effect, the two cross-ply layups (902/02)s and (90/0)2s with 0.5 mm respectively 0.25 mm layer thickness are tested in a specialised VHCF four-point bending test system. For every layup a test series with up to seven comparable load levels is conducted and evaluated.
Cross-ply layups are favourable because of lower damage interactions, compared to quasi-isotropic layups. They develop transverse cracks in the outer layers and delaminations grow from the crack tips along the ply interface.
The composite material consists of the fatigue-optimised fibre SE2020 from 3B-fibreglas and the epoxy system RIM135. The material characterisation shows a strongly improved fibre-matrix adhesion. Due to the glass fibre and the manufacturing in the RTM process the laminate is of very high quality and highly transparent.
Therefore, transmitted light photography can be used to determine the two important fatigue damage parameters crack density and delamination area ratio. Also, general damage mechanisms and the flexural modulus degradation are assessed to identify the influence of the ply thickness.
Comparing the photographs of both test series, a higher crack density is immediately noticeable for the thin layers. Quantitative data determined from the photographs by a specially developed damage detection software confirms this and shows nearly similar delamination area ratios.
With a finite element model, the influence of the crack density and delamination area ratio on the damage development is calculated. Plotting the surface strain and the stress intensity factor over the damage parameters, the influence of ply thickness is directly visible and the development of a “damage path” and different ratios between the damage parameters can be understood.
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QF marker research
Risto Laakso, Keijo Koski, Aleks Vainionpää, Aslak Siljander
Abstract: Experimental crack growth rates using actual load histories increase confidence level of a structural lifetime in two ways: damage rates can be used as a direct estimation method in defining the remaining useful lives, and in selection of the proper calculation parameters.
Post-test quantitative fractography (QF) together with embedded marking sequences in the fatigue loading provides a way to the experimental determination of crack growth curves. The periodically applied markers leave a trace along the fracture surface of the progressing crack. The identification of the traced markers depends on the load and condition history, and can be done with e.g. an optical microscope or a scanning electron microscope (SEM).
The main research questions are: a) how to define and add marker blocks to a given real usage spectrum of a fighter aircraft, b) the former in terms of effects on both small aluminium specimen and full-scale fatigue tests, c) the former two without affecting to the structural lifetime of the test article, and finally d) the applicability of the QF in analysing small crack growth. The future research question is how to use machine learning to automate the markers’ identification process.
Functionality of marker loads especially in relatively short and slow fatigue crack growth (FCG) area is of primary interest. If every loading spectrum contains natural or inserted markers, distinguishability might become a problem during the early stages of the FCG. The marker loading can also appear differently on the surface as the crack growth progresses.
The work started with a literature review, followed by a pre-planning of the marker loads. The pre-chosen marker blocks were calculated with crack growth analysis to see the effects resulted on the structural life. Specified fatigue tests were arranged to obtain experimental results. The research is ongoing, thus only preliminary, and intermediate results can be presented. Taken the intermediate objectives into account, the differences between the QF results were examined: the hypothesis to use an optical microscope alone, cost-effectively, and appropriately in tracing the crack growth in small crack sizes was tested by comparing findings to SEM usage of this and previous projects.
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A dislocation density-based model for the temperature dependent anomalous behaviors of nickel-based single crystal superalloy
Pin Lu, Xiaochao Jin, Xueling Fan
Abstract: Ni-based single crystal has been used as the critical hot-end components of aeroengine due to the excellent mechanical properties. In addition, the finite element method is widely used in performance assessment of hot-end components in engineering applications. Therefore, it is of great significance to construct a constitutive model that can accurately capture the mechanical response of Ni-based single crystals for simulation analysis.
In this work, a dislocation density-based single crystal plasticity constitutive model was developed to capture the temperature dependent anomalous yield and tension/compression asymmetry behaviour of Ni-based single crystals. Firstly, thermally activated cross-slip mechanism, which was considered as the main inducement of anomalous yield behaviour, was introduced into the hardening model. Secondly, the transformation of dislocation motion mode from shearing to by-passing was described using a temperature dependent function. Thirdly, non-Schmid stress tensors were introduced into the constitutive model to describe the tension/compression asymmetric phenomenon. The model considers the contributions of various strengthening mechanisms, including solid solution, precipitates and base metal. Furthermore, the evolution of microstructural features (such as dislocation density, etc.) and contribution of each mechanism to yield stress with increasing temperature were further analyzed. The model has been implemented via crystal plasticity framework and can accurately predict the temperature dependent anomalous characteristics of yield stress of Ni-based single crystal. This work provides a basis for accurately describing and predicting the cyclic loading behaviour of Ni-based single crystals.
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On the mechanism of cyclic crack propagation in AA2024 T3 alloy.
Milan Krkoska, Ligeia Paletti, Rene Alderliesten
Abstract: Fatigue cracks are known to be strongly affected by the type and parameters of the applied loading. The effect of the constant amplitude (CA) and underload (UL) loading was previously investigated on AA2024-T3 alloy in a large volume of research. It was observed that the morphologies of formed fracture surface striations, UL markers (surface ridges) and fissures were directly affected by the magnitude of applied loading cycles, and they are linked to the resulting crack depths and orientations of tilted local fracture surface planes. Despite the large amount of experimental evidences, a comprehensive understanding of the mechanisms governing the observed crack propagation behavior is still to be identified.
Based on an experimental investigation using the so-called crack tip freezing approach, a novel mechanistic model, concerning the cyclic crack tip growth in ductile alloy and the formation of fracture surface features in the wake of progressing crack tip, is presented in this publication. The model covers the formation and the evolution of sharp V-shaped and blunted U-shaped profiles along the crack front, as a function of cyclic loading, and in relation to the crack depth and the local fracture plane tilt. The most significant aspect of the proposed mechanistic model is the incorporation of cyclic plastic deformation of the newly formed crack tip flanks as well as the deformation of previously formed fracture surface features positioned in the wake of the active crack tip. This cyclic, out of plane, deformation explains how the surface striations are shaped into convex profiles on both directly mating surfaces, which could not be previously explained. The developed model also describes the nucleation of the fracture surface fissures at the basis of newly formed crack tip flanks during the application of tensile load excursion and their further growth in the wake of the crack tip via the cyclic plastic deformation. The formation of the asymmetrically shaped crack tip profiles, plastic deformation in the crack wake and the formation of fracture surface features (typical of larger crack depths) are attributed in the model to the deviation of the propagating crack from the general fracture plane.
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