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Über dieses Buch

This book gathers papers presented at the 36th conference and 30th Symposium of the International Committee on Aeronautical Fatigue and Structural integrity. Focusing on the main theme of “Structural Integrity in the Age of Additive Manufacturing”, the chapters cover different aspects concerning research, developments and challenges in this field, offering a timely reference guide to designers, regulators, manufacturer, and both researchers and professionals of the broad aerospace community.



Additive Manufacturing


Fatigue Characteristic of Linear Friction Welded Ti-6Al-4V Joints

A Blisk is the integrated part of rotating blades and a disk, which is recently being adopted for fan and compressor modules of jet engines for the purpose of the weight reduction and the performance improvement. The blisk is generally manufactured by milling from the large forged material. Therefore large amount of material is wasted as cutting chips. Linear friction welding (LFW) is expected to save the wasted material, which is a kind of the solid state joining. The purpose of this study was to investigate the reduction of the fatigue strength due to LFW process. The fatigue test specimens of LFW joints and the base material were made from the same forged material to minimize the effect of scatter. Fatigue tests and analysis of covariance were conducted, and it was made clear that the fatigue strength of as-welded joint was slightly lower than that of base material, and that that of post weld heat treated (PWHTed) LFW joint was the same as that of base material. Since it was thought that the fatigue strength reduction of as-welded LFW joint was due to the effect of residual stress, the residual stress was measured before and after PWHT by Center Hole Drilling method. It was confirmed that the residual stress of as-welded LFW joint remained approximately 3 mm from the weld line where the fracture occurred in the fatigue tests, and that the residual stress of PWHTed LFW joint was reduced drastically. As a conclusion, it was made clear that the fatigue strength of PWHTed LFW joint was equivalent to that of base material with the statistical data.

Hiroshi Kuroki, Yukihiro Kondo, Tsukasa Wakabayashi, Kenji Nakamura, Kikuo Takamatsu, Koji Nezaki, Mitsuyoshi Tsunori

Fatigue in Additive Manufactured Aircraft: The Long Way to Make It Fly

Manufacturing of metallic products is currently facing a historic revolution driven by a group of innovative technologies clustered under the name of Additive Manufacturing (AM) or, more colloquially, 3D-printing.As opposed to the conventional subtractive manufacturing methodologies, AM is able to create net-shaped products by means of the addition, layer by layer, of the needed material.The biggest benefit of the AM in respect to the conventional manufacturing technologies is the substantial freedom the technology gives to the product designers in frame of the product development.Due to the relaxed geometrical constraints, complex parts characterized by optimized shapes can be realized.This can be translated in development of lighter and more cost-efficient parts and assemblies hitting two of the most important targets of the aeronautical industry. Furthermore, as a consequence of the significant weight benefits, the application of topologically optimized AM aircraft structures can provide with an important contribution in reducing the aircrafts fuel consumption and the relative emissions of CO2 & NOX gases, in line with the environmental targets of current aviation research programs.Despite the high investments, the AM still finds a limited applicability for the manufacturing of aircraft components.This is also due the challenges relative to the fulfillment of the Fatigue and Damage Tolerance (F&DT) requirements, particularly sensitive when they have to cope with innovative materials and technologies.This paper presents the state of art of the AM technology in one of the leading companies specialized in manufacturing of large aircraft components and offers a potential way forward toward the wider exploitation of AM technology potentials.

Ivan Meneghin, Goran Ivetic, Matthias Stiller, Gianluca Molinari, Vjola Ristori, Sara Della Ratta, François Dumont

High Cycle Fatigue and Fatigue Crack Growth Rate in Additive Manufactured Titanium Alloys

The Wire + Arc Additive Manufacture (WAAM) process can produce large metal parts in the metre scale, at much higher deposition rate and more efficient material usage compared to the powder bed fusion additive manufacturing (AM) processes. WAAM process also offers lead time reduction and much lower buy-to-fly ratio compared to traditional process methods, e.g. forgings. Research is much needed in the areas of fatigue and fracture performance for qualification and certification of additive manufactured aircraft components.In this study, specimens made of WAAM Ti-6Al-4V alloy were tested and analysed focusing on two key areas of structural integrity and durability:(1) High cycle fatigue and effect of defects: crack initiation at porosity defects was investigated via fatigue and interrupted fatigue-tomography testing performed on specimens with porosity defects purposely embedded in the specimen gauge section. Key findings are as follows. Presence of porosity did not affect the tensile strengths, however both ductility and fatigue strength were significantly reduced. Fatigue life could not be correlated by the applied stress, e.g. in terms of the S-N curves, owing to the different pore sizes. Using the fracture mechanics approach and Murakami’s stress intensity factor equation for pores, good correlation was found between the fatigue life and stress intensity factor range of the crack initiating defects. Predictive methods for fatigue strength reduction were developed taking account of the defect size, location, and distribution.(2) Fatigue crack growth rate: effect of heterogeneous microstructure was investigated via two different material deposition methods and testing two crack orientations. Fatigue crack growth rates were measured for damage tolerance design considerations. Unique microstructure features and their effect on the property anisotropy are discussed.

Xiang Zhang, Abdul Khadar Syed, Romali Biswal, Filomeno Martina, Jialuo Ding, Stewart Williams

Strain Controlled Fatigue Testing of Additive Manufactured Titanium Alloy Ti-6Al-4V

This paper describes strain controlled fatigue testing of a titanium Ti-6Al-4V alloy, additive manufactured by “electron beam melting” (EBM). The EBM material is manufactured in two conditions; with no post-manufacture heat treatment (“As-Built”) and after a hot isostatic pressing (HIP) treatment. The EBM HIP treatment condition is manufactured in three build orientations; vertical, horizontal and at 45°. The fatigue test results for these EBM material conditions are compared with those for similar titanium Ti-6Al-4V alloy powder, manufactured by powder metallurgy hot-isostatic pressing (PM HIP), and for similar titanium Ti-6Al-4V alloy manufactured by traditional wrought mill into bar and sheet material.The strain-life fatigue damage model and fatigue characterisation method used to fit fatigue test results from traditional manufacturing methods (wrought and PM HIP) appears to be applicable to the additive layer manufacturing method (EBM) for this titanium Ti-6Al-4V alloy material.The EBM As-Built and HIP conditions in the low-cycle region all show similar fatigue performance. This is expected given their similarity in tensile strength. The effect of the HIP on the EBM additive manufactured material is seen in the high-cycle region with much better fatigue performance. This is expected as the HIP treatment reduces porosity in the material and improves the fatigue life. The three EBM HIP build orientations all show very similar fatigue performance, though the vertical has slightly longer lives than the corresponding horizontal and 45° build orientations. It is not possible to identify whether these slightly longer lives are because of a build orientation difference, a build-to-build difference, or an effect of powder recycling.In conclusion, fatigue tests on additive manufactured material, including both manufacturing process and any post manufacturing treatment, is considered essential because the fatigue performance of additive manufactured material cannot be inferred from tensile tests or from comparable wrought material.

Rob Plaskitt, Andrew Halfpenny, Michelle Hill

The Optimization and Design of Complicated-Surface Panel Based on Automate Fiber Placement

Aiming at the special configurations of the automated fiber placement composite panel and the complex loads, the applicable analysis criteria and failure criteria are screened and perfected. Iso-parametric shell element is used to simulate skin and stiffener. A specific interface element is used to simulate between skin and stiffener. During the analysis process of the buckling, post-buckling and failure of the structure under complex loads, the modified Hashin criterion is used to simulate the longitudinal and transverse failure of the material. The tensile and shear failure criteria are used to simulate the interface between skin and stiffener. With the increase of the number and height of stiffeners, the buckling and post-buckling capacity of double-curvature stiffened panel increases. But the buckling and post-buckling capacity of longitudinal and transverse stiffeners are very close.

Tieliang Zhang, Liyang Liu, Hao Cui

Advanced Materials and Innovative Structural Concepts


A Modeling Approach for the Fatigue Behavior of Laser Drilled Micro Perforated Structural Panels

With hybrid laminar flow control the drag can be reduced for airfoils. This is done by boundary layer suction through millions of small holes in the skin panels. The goal of this paper is to assess the impact on fatigue properties and to describe a modeling approach to account for the laser drilled holes in the skin panels. Two fatigue test programs were performed to obtain inputs to model the fatigue behavior of micro perforated titanium panels. From the test data several SN curves were generated. The all the SN curves have a similar slope as the basic material. This allows for a factorization approach to model the fatigue behavior. Three factors are applied to the basic material properties. The first factor takes into account the hole geometry, which be captured analytically with a stress concentration factor. The pitch between the holes is large enough that there is no interference for fatigue initiation. Secondly multiple site damage was visually observed by multiple crack plateaus on the crack surface. The crack jumps between the rows of holes. This can be seen as shift in the SN curve which is in line with the knock down factor found by an analytical approach. And finally a technology factor to account for the manufacturing process was obtained. The manufacturing process for the coupons of the two test programs was different, resulting in a different technology factor. Therefore setting specific requirements for the manufacturing process can reduce the impact on the fatigue properties. By multiplying the basic material properties with these three factors according to Eq. (5) it is possible to do standard fatigue assessments in commonly used fatigue tools and with detailed spectra.

Dort Daandels, Stefan Riekehr, Nikolai Kashaev, Jon Mardaras, Sammy Zein El Dine, Christian Heck

Assessment of Fatigue Behavior of Advanced Aluminum Alloys Under Complex Variable-Amplitude Loading

The Federal Aviation Administration (FAA), Arconic, and Embraer have partnered in an effort to evaluate several emerging metallic structures technologies (EMSTs) through full-scale tests and analyses. The focus of this paper is a supplemental test program that is being carried out by the FAA and Arconic in support of the full-scale panel tests. The full-scale fixture is capable of executing complex variable amplitude spectrum loading; however, it was not considered practical for this program to run a full-scale fatigue or damage-tolerance test under such conditions. Instead, an equivalency approach using M(T) fatigue crack growth specimens was conducted to determine an equivalent constant amplitude loading in the axial direction that can be used in a test to accurately represent the complex flight loads. It was also of interest to determine whether to include the once-per-flight compressive load simulating landing in the spectrum crack growth M(T) testing. The special test method employed, results of the equivalency testing, and a comparison to the full-scale panel test results are covered.

Kevin Stonaker, David Stanley, John G. Bakuckas, Mike Kulak, Po-Yu Chang, Gongyao Wang, Mark Freisthler

Damage Mechanisms and Mechanical Properties of Directly Bonded CFRTP and Aluminium with Nano-Structured Surface

The present work studies the cohesive behaviour of a previously proposed novel direct bonding method for dissimilar bonding between a carbon fibre reinforced thermoplastic (CFRTP) and aluminium. A nanostructure is manufactured on the aluminium surface and is directly bonded to the CFRTP by applying heat and pressure. Double cantilever beam (DCB) testing is carried out to evaluate the bonding properties and the initial results of a method for directly measuring the traction-separation behaviour from experiments is presented. The nanostructure is observed to improve the bonding properties significantly compared to two other considered bonding cases. Furthermore, the measured traction-separation behaviour is seen to be difference for each case. Nevertheless, the applied calculation method shows some challenges related to thermal stresses and plastic deformation that should to be taken into account in future studies.

Kristine Munk Jespersen, Hikaru Abe, Hiroki Ota, Kei Saito, Keita Wada, Atsushi Hosoi, Hiroyuki Kawada

Interaction Between High- and Low-Cycle Thermo-Mechanical Fatigue Crack Propagation Around Cooling Hole in a Ni-Based Superalloy

Cooling hole area introduced in gas turbine blades may be one of susceptible areas to fatigue failures, when gas turbine systems are subjected to frequent load change. Here an interaction between high- and low-cycle thermo-mechanical fatigue failure is an key issue to be concerned. In order to get basic understandings on the structural reliability under such a condition, a new testing system has been developed in this work. By means of the test system the propagation behavior of the small crack nucleated from a simulated cooling hole in a directionally solidified Ni-base superalloy was studied under the artificial condition in which the high-cycle thermo-mechanical fatigue (TMF) loading was superimposed on the low-cycle TMF loading. The experimental works demonstrated that the role of the high-cycle thermal stress cycle resulting from non-stationary response of the structure significantly interacted with the stationary low cycle TMF loading.

Masakazu Okazaki, Yuuki Yonaguni

Ply Curving Termination to Suppress Delamination in Composite Ply Drop-Off

Composite tapered sections with ply drop-off are prone to delamination due to stress concentration at the edge of the terminated 0° plies. Currently quite gradual taper angle is used in aircraft composites to suppress delamination, which leads to weight increase. Structural modification that can suppress the stress concentration is necessary to realize further-optimized lightweight structures with steeper taper angles. This study proposes a new approach called “ply curving termination (PCT)” in which local fiber orientation is modified at the edge of the terminated 0° plies. The curved area has significantly low stiffness and thus the stress concentration at the ply edge is suppressed. This paper reports the numerical analysis and experiment to validate the effectiveness of this new concept. In addition, suppression of edge delamination by PCT is also demonstrated.

Shu Minakuchi, Nobuo Takeda

Studies on the Fatigue Damage Behavior of Active Jet Engine Chevron

The combination of theoretical research and finite element simulation is used in this work to study the fatigue damage behavior of active jet engine chevron. The main research contents are as follows:Firstly, the thermo-mechanical constitutive model is developed to describe the behavior of shape memory alloys considering the damage. Based on the continuum damage mechanics theory, assumed that when the damage happens, the mass density, the elastic compliance tensor, the thermal expansion coefficient tensor, the specific heat, the specific entropy and the maximum phase change strain are all caused damage. The model is started from the second law of thermodynamics, a detailed procedure for the estimation of the stress-strain relationship is presented while the plastic strain caused by phase change is ignored. Secondly, numerical results of the stress-strain relationship for the shape memory alloys under different damage extent are discussed.Finally, a finite element model of the jet engine chevron is established, the fatigue damage behaviors of active jet engine chevron are studied under the cycle loadings, and the tip deflection of the jet engine chevron when the fatigue damage happened is discussed.

Bingfei Liu, Shangyang Jin, Shaozhe Dong, Zhenyu Feng

Airworthiness and Other Considerations


An Ultrafast Crack Growth Lifing Model to Support Digital Twin, Virtual Testing, and Probabilistic Damage Tolerance Applications

New aeronautical technologies like the Airframe Digital Twin, Virtual Fatigue Testing, and Probabilistic Damage Tolerance Analysis require a very large number of crack growth evaluations with a comprehensive number of random variables in order to accurately predict the fatigue life, the structural risk, or the remaining useful life of a structure. Current state-of-the-art crack growth methodologies and probabilistic methods do not make these new technologies possible due to limitations on computational speed, number of random variables, and statistical tools. In this work, a new computational strategy is developed and demonstrated such that several random variables directly affecting the crack growth analysis can be considered. This approach provides the opportunity for a more comprehensive and accurate digital twin evaluation, virtual testing prediction, and risk assessment, hence, improving aircraft design, safety, and reliability.Under Federal Aviation Administration (FAA) Funding, This methodology focused on the development of an ultrafast numerical crack growth algorithm that consists of: (a) a constant amplitude equivalent stress derived from a variable amplitude loading spectrum, and (b) an adaptive step-size Runge-Kutta ordinary differential equation (ODE) solver.Several examples with the Airframe Digital Twin, Virtual Fatigue Testing, and Probabilistic Damage Tolerance applications will be demonstrated using through, corner, and surface cracks at a hole under representative loading spectra. The crack size versus cycles results from this new approach will be compared against results obtained from commercial lifing software codes. All results to date indicate the comparison is within a few percent. The probabilistic crack growth analysis has been parallelized using OpenMP in order to fully utilize multi-core computers. This approach provides a more comprehensive and accurate risk assessment, hence, improving aircraft safety and reliability.

Juan Ocampo, Harry Millwater, Nathan Crosby, Beth Gamble, Christopher Hurst, Michael Reyer, Sohrob Mottaghi, Marv Nuss

Analytical and Numerical Investigation of the Effect of Secondary Bending in Hard-Point Joints

Typical fuselage structures consist of several structural details, such as skins, frames, stringers, doublers and tear straps, joined together by means of mechanical fasteners. Upon cabin pressurization, membrane stresses are induced in the skin sheets every flight. Fastened attachments include inherent eccentricity with respect to the load path, which induces out of plane deflections of the skin sheets. This is referred to as ‘secondary bending’, and it is considered as a side effect of tensile membrane loads acting on the skin. In this study, the effect of the secondary bending on hard-point attachment is investigated analytically and numerically, and the relation σbending/σtenslie is obtained for different geometrical combinations and different hard-point scenarios (with and without skin cutout). Good agreement between analytical and numerical results is reported. The maximum ratio obtained numerically for typical fuselage configurations was derived as σbending/σtenslie = 0.20, which is significantly lower than that obtained for lap joint attachments (that can reach up to σbending/σtenslie = 3.0). An example of an antenna installation is presented, showing significant decrease in the fatigue and crack growth lives due to the induced secondary bending effect.

Yuval Freed, Lior Sagi Machnes, Orel Magidish

Demonstration of an Airframe Digital Twin Framework Using a CF-188 Full-Scale Component Test

The airframe digital twin (ADT) framework is a potential game-changing fleet management concept recently proposed by the United States Air Force to allow proactive and cost-effective decisions on an individual aircraft basis. The National Research Council of Canada is currently demonstrating the ADT framework using a CF–188 full-scale component test to assess the adaptability of this approach for Royal Canadian Air Force (RCAF) fleets. An in-house analysis tool is being developed to perform quantitative risk assessment (QRA) based on the Bayesian inference method using individual aircraft tracking and non-destructive inspection data. The modular components of the ADT tool, currently being validated, include load and crack size distribution updating, material initial discontinuity state, residual stress effects, load transfer functions, crack tip stress intensity factor calculations, and crack growth predictions. Test cases analysed to verify and validate these modules showed the benefits of the Bayesian updating approach for performing QRA with inputs that are initially scarce and become less uncertain throughout the service life of the aircraft. Short-term benefits expected from the application of the ADT approach in the RCAF fleet management include a better use of the IAT data and an improvement in fatigue life estimation. In the longer term, a higher return on investment is foreseen in terms of improved life cycle management, increased fleet availability, and reduced total fleet ownership costs.

Guillaume Renaud, Min Liao, Yan Bombardier

Development of Efficient High-Fidelity Solutions for Virtual Fatigue Testing

Virtual Fatigue Testing (VFT) can be defined as the high-accuracy simulation of the behaviour of high-fidelity ‘digital twins’ representing physical test articles in their test environment. The final aim of virtual fatigue testing is to provide an alternative to physical fatigue tests as the main source of certification evidence, as requested today by most civil and military airworthiness regulations, thus enabling the so-called ‘certification by analysis’.This paper presents the work developed by Airbus towards the practical implementation of VFT capabilities for the assessment of military aircraft. A key factor for the eventual success of VFT is the ability to select and incorporate the latest simulation techniques in this field while keeping a reasonable computational cost. The selected methods need to cover all the phases of a standard fatigue test: crack initiation, crack growth and residual strength. The scope of VFT will include also the systematic evaluation of the statistical uncertainties replacing the traditional use of scatter factors.Finally, the overall quality of VFT as certification evidence is addressed, because the fulfilment of the eventual requirements for the acceptance of virtual tests by the Airworthiness Authorities is an integral part of the development. Therefore, the criteria for verification, validation and final certification being created for this purpose are also discussed.

Javier Gomez-Escalonilla, Diego Garijo, Oscar Valencia, Ismael Rivero

Effective Durability and Damage Tolerance Training: New Methods for Modern Learners

In the aeronautical industry, development of fundamental skills in durability and damage tolerance is critically important for most engineers engaged in structural design and analysis tasks. These skills are not usually well covered in most U.S. university undergraduate curricula, and individual coaching or on-the-job training are simply not efficient means of acquiring these skills at a basic level. An additional consideration is that the traditional classroom instruction formats are losing some of their effectiveness, as modern learners entering the workforce today have grown accustomed to a more hands-on style of learning. As a pilot program in transforming the Boeing Commercial Airplanes (BCA) Structures Engineering New Hire (or foundational) training, we have recently rebuilt the introductory durability and damage tolerance (DaDT) portion of this curriculum, originally developed nearly 20 years ago, with a more learner-centric focus. The introductory DaDT syllabus covers the fundamentals of durability, fail-safety and damage tolerance, and how these principles are used at BCA. Boeing has thousands of engineers engaged in work that is pertinent to the structural integrity of our commercial airplanes. Many of these engineers go on to take more in-depth training in which they learn how to apply our proprietary DaDT methods. We have recently taken a new approach for this introductory course, which embraces modern learning methods as a means to increase student engagement and learning and knowledge/skill retention. In this paper we provide an overview of this endeavor and we share our learning architecture model, the perceived benefits, testimonials from learners, and our plans for the future.

Brandon D. Chapman

Fatigue Considerations in the Development and Implementation of Mechanical Joining Processes for Commercial Airplane Structures

The durability of commercial airplane structures is strongly influenced by build quality and the extent to which assembly processes can be controlled such that the fatigue quality is consistent with that assumed in the design throughout the production life of airplane programs. New developments in assembly technology and the continual quest for manufacturing cost reductions, as well as rising production rate pressures are translating into new, often non-traditional parts and build processes being applied to commercial airplane products. This technological evolution can only take place by exercising close coordination between the structural and manufacturing engineering functions in planning, evaluating, and bringing these parts and processes safely into production. Described in this paper are some of the new processes being used at Boeing Commercial Airplanes to produce metallic and hybrid (composite + metal) assemblies for large commercial transports, viewed from a broad structural engineering perspective. The discussion focuses on two specific mechanical joining technology thrusts: (1) One-up assembly (OUA) and process automation, and (2) assembly using pre-drilled holes at the part fabrication (detail) level. A number of case studies are outlined and considerations such as process selection, control, and qualification, and fatigue characterization are highlighted. This includes a discussion of the trade-offs in fatigue capability between traditional and new methods, including some quality issues that can arise with the new approaches.

Robert Jochum, Antonio Rufin, Tanni Sisco, Frederick Swanstrom

Rapid Calculation of Safe Acceleration Values for Aircraft Structures Under Flight Test

During flight test programmes structural response data may only be available in terms of accelerations, but relating these accelerations to stresses can be difficult without extensive analysis. Presented here is a method to quickly and accurately generate allowable acceleration levels to prevent fatigue failure. The method applies to lightly damped structures primarily excited at their fundamental mode such as antennae, radomes, and panels, but may be able to be conservatively extended if multiple modes exist.The method relies on the structure primarily responding at its fundamental mode, and therefore that frequency can be used for the upwards crossing rate and also used to govern the relationship between acceleration and displacement. Furthermore, at the fundamental mode the relationship between displacement and detail stress can often be estimated based upon quasi-static considerations. Accurate understanding of the damping in the system is not required. Together these allows simplifications to be made to relate accelerations with stress and then fatigue damage.Accuracy of the results is evaluated and quantified, and the situations where acceptable accuracy achieved is defined. Finally, a nomogram is provided to aid the rapid calculation of allowable acceleration levels.

Stephen Dosman, Jonathan Gorman

Reliability Approach Applied on Fatigue Safety Factors for Fleet Monitoring

French armament procurement agency (DGA) is currently implementing a reliability approach in order to check aircraft safety level. The goal of the method is to carry out a quantitative study about the structural failure based on random phenomenon. Physical randomness is considered in the aviation regulation with the use of scatter factors. Dealing with safe-life substantiated aircraft, DGA, in charge of airworthiness, allows the reduction from 5 to 3 of the fatigue scatter factor for the monitored fleets with a safety level at least preserved. Two aspects are analyzed with the reliability method: The first application evaluates the safety level induced by different monitoring systems. The second one quantifies the safety level of aircraft without monitoring system compared with monitored ones in order to evaluate the reduce scatter factor.

Vincent Montlahuc

Research on the Airworthiness Compliance Strategy of Composite Structure

Composites have been widely used in the main structure recently, because the automated production process has more stable production quality, more efficient production efficiency, lower production costs, so the automation technology is highly applied. In this paper, the automation processes of various current and future applications are analyzed, and the challenges encountered in the design are put forward according to the characteristics of the processes. Based on the research of the requirements of composite material, process, its structure design, analysis, manufacturing technologies, verification process, maintenance and repair, etc. give the airworthiness compliance method and test planning suggestions to provide technical support for primary composite airframe structure application in the civil aircraft.

Weiping Li, Xiaoling Zheng

Risks of Initial Assumptions in Fatigue and Damage Tolerance of Small Aircraft Development Programs

Structural Integrity Programs as part of the development of new aircraft, a major change to type design, a continued airworthiness program, or the development of a supplemental structural inspection document; require a large set of initial assumptions, particularly in the area of fatigue and damage tolerance. This review paper outlines assumptions, the associated risks, and methods of how to manage these risks. Airplane or Rotorcraft certified under the requirements of Part 23 and Part 27 respectively, are considered as “small”. The selected certification basis defines the fatigue evaluation requirement, where different concepts are applicable depending on the type of component, the design, and the material. The structural development process involves the steps of mission and usage assumptions, spectrum development for components including principal structural elements, fatigue load derivation, material selection, and material data derivation as part of the building block testing approach. This process includes development, certification, and qualification tests to cover conventional designs and advanced manufacturing technologies. Damage tolerance analyses require software tools and adequate libraries of Stress Intensity Factors. The selection of scatter factors and life improvement factors for interference fit fasteners, cold worked holes, or shot peened structure as well as the definition of initial and detectable crack sizes with their probability of detection are important parameters. Every step involves assumptions and simplifications with associated risks, that are often covered by conservatism and defined by judgement, however, when quantification becomes inevitable, probabilistic approaches should be employed.

Dejan Romančuk, Juan Ocampo

Russian Practice to Provide Safe Operation of Airplane Structures with Long-Term Operation

The article presents some general approach to provide safe operation of aging aircraft structures including those with long-term operation (Fig. 1). Data on increased service life values of aged aircraft in Russia are given. Methods to detect and eliminate the structural corrosion damages are discussed. Statistical data on distribution of typical defects in civil aircraft structures are given including some results on statistical analyses of the duration of corrosion depth increase in wing and fuselage skins of IL-86 wide-body passenger aircraft. Authors outline the approaches to prevent the failure of aircraft structures due to widespread fatigue damage with provided statistics on the multiple site fatigue damages in panel joint of the upper wing surface of Il-62 M civil airplane that were identified while full-scale tests of aircraft and obtained from operation. In addition, the results of experimental research on degradation of strength, fatigue, and crack growth resistance of long-term-operated airframes are given. Fig. 1. Compared operation time of Russian airplanes to their design goals. Dotted lines correspond to individually extended service life fleet leaders

Boris G. Nesterenko, Grigory I. Nesterenko, Victor V. Konovalov, Vitaly Ya. Senik

Smarter Testing Through Simulation for Efficient Design and Attainment of Regulatory Compliance

Passenger safety is by far the most important consideration in the development and operation of commercial aircraft. How does Boeing ensure that the structure and systems on its aircraft meet regulatory requirements? A rigorous building block approach verifies and validates analysis by tests. Best practices in tasks such as material characterization, finite element modeling and simulation methods are documented, standardized and applied to produce high quality predictive assessments. These assessments positively impact design and enable well-planned and informed smart testing. Smart testing through simulation maximizes the benefit of necessary tests, augments understanding of performance within and beyond the envelope of test data and minimizes unplanned tests in attaining certification. This paper highlights a few examples of how this approach, referred to as “smarter testing,” has been applied at Boeing in recent years to reduce the cost and flow of structural validation required to bring new materials, technologies and non-traditional structural architectures to market.

Steven A. Chisholm, Jack F. Castro, Brandon D. Chapman, Kazbek Z. Karayev, Andrea J. Gunther, Mohammed H. Kabir

Widespread Fatigue Damage Evaluation for Multiple Elements Based on Probabilistic Approach

In order to obtain a Type Certificate for civil transport category aircraft the Applicant must to comply with some design rules - airworthiness requirements. Among those, the one associated with structural fatigue (§25.571), whose main objective is related to prevent a catastrophic event, from structural damage, during the operational life of the aircraft. That section requires special attention for Widespread Fatigue Damage (WFD). For that, the design approval holder (DAH) must establish a Limit of Validity (LOV) of the engineering data that supports the structural maintenance program. Up to the LOV, DAH must demonstrate that the aircraft will be free from WFD. A Widespread Fatigue condition can be originated from: Multiple Site Damage, Multiple Element Damage or a combination of both. The objective of this work is to propose a probabilistic approach to define maintenance actions to prevent widespread fatigue damage condition from Multiple Element Damage up to the LOV.

Fabiano Hernandes

Fatigue Crack Growth and Life Prediction Methods


A Framework to Implement Probabilistic Fatigue Design of Safe-Life Components

Safety-critical components, such as aircraft landing gear, are designed using the ‘safe-life’ fatigue analysis process. Variability exists within materials data, loads data and component dimensions and is currently mitigated using safety factors. Probabilistic approaches to safe-life fatigue design have been proposed to better represent this variability. However, challenges currently exist that prevent the wider utilisation of a probabilistic approach. This paper presents a framework that aims to overcome these challenges. The statistical characterisation, probabilistic, surrogate modelling and sensitivity analysis methods required to implement the framework are introduced. Finally, a discussion of how recent advances within aerospace fatigue design, such as ‘big-data’, can be used to support a probabilistic framework is presented.

Joshua Hoole, Pia Sartor, Julian Booker, Jonathan Cooper, Xenofon V. Gogouvitis, Amine Ghouali, R. Kyle Schmidt

A Multiaxial Fatigue Damage Model for Isotropic Materials

This paper presents a novel damage mechanics based failure model enabling the prediction of low cycle fatigue life and residual strength of isotropic structures under multiaxial loading. The approach herein proposed does not discretize every load cycle but instead takes an envelope loading whereby the numerical load remains constant at a maximum load level and the number of cycles is obtained from a given elapsed time defined within a pseudo-time framework. The proposed formulation is based on the smeared cracking approach accounting for damage propagation due to static and fatigue loadings, where the static component is based on the Von-Mises yield criterion and Prandtl-Reuss stress flow rule; whereas the crack propagation in cyclic loading component is based on the Paris-law. Furthermore, the formulation combines damage mechanics and fracture mechanics within a unified approach enabling the control of the energy dissipated in each loading cycle.

Mauricio V. Donadon, Mariano A. Arbelo, Paulo Rizzi, Carlos V. Montestruque, Lucas Amaro, Saullo Castro, Marcos Shiino

A Specimen to Evaluate Susceptibility of Aluminium Alloys to L-S Crack Deviation

The current aircraft structures tend to increase the proportion of integral structure parts. One main advantage is the cost, through a reduction of the assembly complexity. The fatigue behaviour of integral structures is also improved due to a reduction in potential initiation sites at joints or rivets, and the stiffness is considered better. The design of new parts requires, however, an analysis of the damage tolerance behaviour (van der Veen et al. 2016). The present study focuses on the fatigue crack propagation of such structures. Complex parts are commonly machined in aluminium thick plates. Several studies demonstrate that cracks may deviate in unstiffened L-S crack configurations for standard alloys (Joyce et al. 2016; Sinclair and Gregson 1997). T-stresses and mixed-mode loads are also affecting the crack deviation behaviour (Llopart et al. 2006). The L-S CT specimen is shown sufficient to differentiate some variations, e.g. variations between three positions through-thickness. However, the less deviating alloys are difficult to distinguish. A new asymmetric four-stiffener specimen (WEND) geometry is proposed as an alternative test. Experimentally, in order to demonstrate the advantages of the WEND specimen, 7010-T76, 2139-T8 and 2050-T8 plates are characterized. The same alloys are characterized using CT specimens. The WEND specimen is a lab-scale test closer to real structures than a CT test. In its first use, it allows to compare an alloy behaviour with a targeted lifetime vs crack path. In its second on-going use, it enables the comparison with Finite Element Modelling for a better crack path and lifetime prediction.

Erembert Nizery, Jean-Christophe Ehrström, Guillaume Delgrange, Bruno Wusyk

A Numerical Approach to the Disbonding Mechanism of Adhesive Joints

The need for lightweight structures in aeronautics is leading to a strong interest in adhesively bonded joints. Incomplete knowledge of their fatigue behaviour is a major obstacle to their application. At present, the prediction of the disbonding growth is yet an open question. This work aims to develope a numerical model for the computation of the disbonding growth in an adhesive joint. The scope is calculating the energy release under quasi-static conditions in order to relate it to the fatigue disbond growth through the existing analytical models. A finite element model for the prediction of disbond growth under quasi-static loading has been implemented in Abaqus, by introducing a cohesive zone model which is able to capture the process zone around the crack tip and to enforce an energy-based failure criterion. The model, which had originally been developed for double cantilever beam specimens under mode I, was extended to mode II loading.Numerical simulations are validated by comparison with experimental results on double cantilever beam coupons in mode I and with literature results on end notched flexure coupons in mode II conditions. The results from tests and simulations are in accordance with each other.The presented model is a suitable option for the estimation of fracture mechanics parameters in cases in which complex geometry and loads prevent the application of analytical theories.

Nicola Zavatta, Enrico Troiani

An Engineering Calculation Method of Probability Distribution of Crack Initiation Life for Widespread Fatigue Damage

An engineering calculation method of predicting probability distribution for crack initiation life of structures susceptible to widespread fatigue damage is proposed. Through a study of the crack initiation mechanism, the incident that crack initiation life of multiple detail structure taking a certain value is transformed into the intersection of three independent incidents. The probability of occurrence of the former incident is the product of the probability of occurrence of the latter three independent incidents, which is derived from the probability distribution function of crack initiation life of single detail structure. Thus, the detailed formulae of probability density functions of initiation lives for cracks appearing in turn in multiple detail structure are obtained. Making use of these detailed formulae, the calculation formula of median rank of initiation life is derived. It is only related to failure order and total number of details. Through the value of median rank, the initiation life with reliability of 50% can be achieved. Thus, an engineering calculation formula of probability distribution of crack initiation life is gotten. Crack initiation test of specimens with a single hole and multiple holes were carried out. The model is used to estimate crack initiation lives of the multiple holes notched specimens. The predicted results are in good agreement with the testing results, which show that this model is effective.

Xi Wei, Li Qiang, Shen Peiliang, Yang Gang, Huang Fu, Zhao Jianjun

Assessment of Aircraft Structural Service Life Using Generalized Correction Methodology

At present, there is no common model or approach that can comprehensively cover all stages and influence factors of the general fatigue and crack growth behavior. Classical fatigue models cannot be directly applied to service life prediction under variable amplitude (VA) load spectra without calibration and validation by fatigue tests data. In this research, a generalized correction methodology is proposed, which is divided into life correction rule and crack growth rate (CGR) correction rule, the former is limited to crack initiation models, and the latter is limited to effective block approach (EBA) model, based on the assumption that the uncalibrated classical fatigue models can provide a reasonable prediction for the relative severity of two VA load spectra. The application procedure of generalized correction methodology is detailed, and the way how to select the durability and damage tolerance (DADT) optimal models is provided. Furthermore, sensitivity analyses of model parameters largely affecting the accuracy of life predictions in crack initiation models and EBA models are presented and suggestions on their derivation are given. Finally, case study of some bulkhead representative coupon data under three test spectra at different stress levels is reviewed where the implementation of generalized correction methodology has been comprehensively examined and validated, which is anticipated to provide highly accurate and reliable structural service life assessment for agile fighter aircraft under operational load spectrum.

Hongna Dui, Xiaodong Liu, Jiang Dong, Lixin Zhang

Examination of the KAWAI CLD Method for Fatigue Life Prediction of Composites

The Kawai Constant Life Diagram (CLD) method was examined for fatigue life prediction of composite materials. Static and fatigue tests were carried out for open-hole coupon specimens made of unidirectional carbon/epoxy tapes for examining the applicability of the KAWAI modified constant life model. The Goodman method was examined as well. The prediction of the Kawai model was slightly conservative for R-ratios ≥ −1 and un-conservative for R-ratio that equals 10. The Goodman CLD highly overestimated fatigue life for purely tension and tension-compression areas and underestimated fatigue life for R-ratios lower than R = −1. A modified Kawai model was suggested to overcome the un-conservatism of the Kawai model at the pure compression zone of the CLD.

Yael Buimovich, Dvir Elmalich

Fatigue Crack Growth Prediction and Verification of Aircraft Fuselage Panels with Multiple Site Damage

A numerical algorithm of fatigue crack growth prediction on aircraft fuselage panels with multiple site damage is discussed in this paper based on Finite Element Method. The objective of this research is to realize the automatic crack propagation and crack growth life calculation. The mixed-mode stress state of structures at crack tip is considered through the published form of effective stress intensity factor, and the incremental fatigue crack growth model is studied to coordinate the propagation between multiple cracks. Maximum tangential stress criterion is applied to determine the crack growth direction at a growth step. Moreover, the new crack tip position in flat plates and curved plates are numerically determined to automatically update the crack tip. Fatigue test of curved panel subjected to inner pressure was conducted by horizontal self-balanced test facility, and the fatigue tests of plates with two internal collinear cracks or with seven collinear cracks under tensile stress were also analyzed to validate the feasibility of algorithm. It is shown that the efficient fatigue crack prediction algorithm is able to predict various crack growth behaviors observed in tests, and the predicted crack propagation path and lives are in a good agreement with test results and data available in the literature.

Shaopu Su, Jianghai Liao, Wendong Zhang, Dengke Dong

Fatigue Life Prediction of CFRP Laminate Under Quasi-Random Loading

The results of fatigue test and fatigue life predictions of CFRP T300/5208 [45/0/-45/90]2s specimens with open holes under loading of quasi-random “TWIST” program with different levels of truncation of large and small loads were considered. It was noted that the fatigue life predictions made using the Palmgren-Miner rule showed an unacceptable accuracy of the results. In order to increase the accuracy of such predictions, two new prediction methods are proposed, formed by using two different non-liner fatigue damage accumulation models. Results of fatigue life predictions for considered specimens with use of new methods are presented. The results of the predictions are compared with the experimental data. Conclusions are made about the accuracy of predictions using the proposed methods.

Vitaly E. Strizhius

Fatigue Life Simulation and Experiment of 2024 Aluminum Joints with Multi-Fasteners Interference-Fit

A three-dimensional finite element model of 2024 aluminum hi-bolted joint was established. Hi-bolts installation and pre-tightening force applied were simulated, then distal alternating load on middle plate of the joint. Fatigue life was predicted with FE-SAFE soft, and verified by experiment. The simulation results show that middle plate is the weakest plate in the plates of the joint. Hoop residual compressive stress increases with raising interference, and decreases gradually from the hole to the outer edge. Fatigue life is longer with interference fit than clearance, and it is benefit for fatigue life improvement with interference from 0.08 to 0.14 mm. While fatigue test results show that failure origins from No. 2 hole of the middle plate. The fatigue life grow longer with interference from 0 to 0.11 mm, then shorten with 0.14 mm interference. The fatigue life is longest with 0.11 mm interference. The suggested interference for engineering is 0.08–0.11 mm for possible defects with 0.14 mm interference.

Qingyun Zhao, Yunliang Wang, Hong Huang, Sirui Cheng, Fenglei Liu

Influence of Heat Treatment on Near-Threshold Fatigue Crack Growth Behavior of High Strength Aluminum Alloy 7010

In this study, aluminum alloy 7010 was subjected to three different ageing treatments i.e., peak ageing (T6), over ageing (T7451) and retrogression and re-ageing (RRA) to study the influence of precipitate microstructure on the fatigue crack growth rate (FCGR) behavior. The microstructural modifications were studied by using TEM to examine the change in size and morphology of the precipitates. The size of the precipitates in the matrix range from 16–20 nm in T7451, 5–6 nm in RRA and 2–3 nm in T6 alloys, respectively. The FCGR tests were performed on standard compact tension (CT) specimens as per ASTM E647 standard in a computer controlled servo-hydraulic test machine with applied stress ratio, R = 0.1 and loading frequency of 10 Hz. The crack growth was measured by adopting compliance technique using a CMOD gauge attached to the CT specimen. The fatigue crack growth rate was higher in T7451 and lowest in RRA treated alloy. The RRA treated alloy showed higher ∆Kth compared to T7451 and T6 treated alloys. The measured ∆Kth was 11.1, 10.3 and 5.7 MPam½ in RRA, T6 and T7451 alloys, respectively. In the near-threshold regime, the RRA treated alloy exhibited nearly 2–3 times reduction in the crack growth rate compared to the T6 alloy. The growth rate in the RRA alloy was one order lower than that of the T7451 condition. The surface roughness of RRA treated alloy was more pronounced. The reduction in FCGR observed in RRA alloy was correlated to partial crack closure due to tortuous crack path and partially due to increased spacing between the matrix precipitates. The reduction in near-threshold FCGR and increase in ∆Kth is expected to benefit the damage tolerant capability of the aircraft structural components under service loads.

M. S. Nandana, Bhat K. Udaya, C. M. Manjunatha

Multiaxial Fatigue Behavior of 30HGSA Steel Under Cyclic Tension-Compression and Reversed Torsion

Many critical mechanical parts in aerospace, automobile and other industries are subjected to complex cyclic loading during their service life. The 30HGSA steel is one of the most commonly used for manufacturing of this highly loaded structures. The 30HGSA has gained a great interest since it exhibits very good strength properties, high hardness, abrasion resistance and contains a trace amount of nickel. Despite its wide applications and superb characteristics the data about material’s behavior under monotonic and combined cyclic in-phase tension-compression and torsion loading is not available in the literature. The paper aims to fill that void by providing a thorough experimental and numerical analysis of the 30HGSA steel. We will examine and compare Gough–Pollard (GP) and Dębski–Gołoś–Dębski failure criteria in the form of limit curves (DGD-LC) and evaluate high-cycle fatigue models. The obtained experimental high-cycle limit curves will be used to make comparison with the above mentioned failure criteria. The results has shown better agreement between experimental data and DGD-LC model than with GP approach.

Daniel Dębski, Krzysztof Gołoś, Marek Dębski, Andrzej Misztela

Novel Methods for Measuring the Mode I and Mixed Modes I/II Interlaminar Fracture Toughnesses of Composite

The present paper devised two novel methods to measure the mode I and mixed modes I/II fracture toughness for composite materials. A double compliance method is proposed for the mode I interlaminar fracture. Only the load and displacement recorded from the test machine are required for the determination of the fracture toughness. It is then used for a DCB test in a cold temperature chamber, where the crack growth length is difficult to get access. Using this method, the finite crack length and fracture toughness can be obtained from the load-displacement relation. In addition, a simple method is proposed to overcome the shortcoming of the current ASTM D6671 by using the mixed modes bending (MMB) specimen. It avoids the measurement of the Young’s moduli from separate tensile tests. The average relative difference between the mixed modes fracture toughness measured from the present method and that from the ASTM D6671 with additional tensile tests to measure the Young’s moduli is within 5%, thus the present paper provides a consistent and simple method for measuring the mixed modes I/II interlaminar fracture toughness. The present methods are simple and accurate, therefore significantly simplifying the procedure for measuring the mixed modes interlaminar fracture toughness.

W. Xu, Z. Z. Guo, Y. Yu, X. J. Zhang

Numerical Investigations on the Three-Dimensional I/II Mixed-Mode Elasto-Plastic Fracture for Through-Thickness Cracked Bodies

Based on three-dimensional (3-d) I/II mixed-mode fracture experiments on an LC4CS aluminium alloy, 3-d I/II mixed-mode elasto-plastic finite element models were established using the commercial software package ANSYS. The coupled effects of the varied degrees of mode-mixing Φ and thickness B on the stress field around the crack tip were analysed, and then the coupled effects of mode-mixing Φ, thickness B, and relative length a/W on the load-crack opening displacement curve were investigated. The results showed that the angle of the maximum tangential stress and the minimum out-of-plane stress constraint factor (Tz) appeared at the same angle with each increment of Φ, and the effects of thickness became weaker with changes to the direction angle of σθθmax and Tzmin with each increment of thickness. The load-crack opening displacement curve was affected by loading angle, relative length a/W, and thickness: the thickness effect was stronger when mode I loading predominated. The load-crack growth length curve can be plotted with reference to the experimental load-crack opening displacement curve, which can be used to predict initiation load in static fracture experiments.

Fang-li Wang, Ming-bo Tong, Shu-wei Bai, Nan Jiang, Chong-min She, Jun-ling Fan

Probabilistic Reliability Assessment of a Component in the Presence of Internal Defects

Structural strength may degrade during the service life of an aircraft due to undetected material defects or accidental damages. Additively manufactured or welded structures are particularly susceptible to fatigue cracking due to stress concentration at the surface and internal material defects. Despite significant benefits provided by the laser-based manufacturing techniques, there is still a lack of understanding how these parts fail under cyclic loading. This study aims to investigate the effect of internal material defects on the high cycle fatigue (HCF) behaviour of Ti-6Al-4V. It is shown that internal fish-eye fatigue fracture is a dominant failure mode if a subsequent surface treatment technique is applied. A probabilistic lifetime assessment framework for predicting the joint durability and its scatter in the HCF regime is developed. A good agreement between the modelling results and experiments is demonstrated.

Fedor Fomin, Nikolai Kashaev

Stress-Intensity Factor Solutions for Tapered Lugs with Oblique Pin Loads

This paper summarizes five years of research into new stress-intensity factor (SIF) solutions for radially advancing cracks at holes in tapered and obliquely loading lugs. To our knowledge, these solutions are the first direct methods capable of analyzing these fracture critical connections. The new SIF solutions rely on weight function (WF) methods, remain tractable for engineering fatigue and fracture analyses, and span a very wide range of applicable geometries (including loading angles between 0 and 90° and independent lug taper angles between 0 and 90°). These solutions support cracks on either side of the pin hole (the “short ligament” or “long ligament” sides). For cracks on long ligaments, users can select the crack location based on maximum principal stresses or Mises stresses. These solutions are verified by a large database of independent high-fidelity three-dimensional finite element analyses (FEA) generated using an automated approach that accelerates the evaluation process while ensuring solution quality. This work also discusses the underlying sensitivity studies used to determine appropriate boundary conditions and material properties. The powerful new methods used here to generate SIFs for tapered lugs—automated FEA generation of tables of stresses in uncracked bodies, interpolation of these stresses in WF formulations, and automated verification using independent benchmark FEA SIF solutions—can easily be extended to build reliable new SIF solutions for other complex combinations of loads and cracked geometries.

James C. Sobotka, Yi-Der Lee, R. Craig McClung, Joseph W. Cardinal

Summary of Recent Round Robin Life Prediction Efforts for Crack Shape and Residual Stress Effects

Two round-robin life prediction efforts were conducted to assess the ability to perform blind life predictions for test data obtained for corner cracks at centered and offset holes in 7075-T651 and 2024-T351 Aluminum alloys. The first round robin effort was conducted as part of a recent AFGROW Crack Growth Life Prediction Software Workshop focused on the ability to accurately predict crack shape evolution, and the second was conducted by the Engineered Residual Stress Initiative (ERSI) Workshop on the effect of split sleeve cold-working. The goal of the ERSI round-robin was to quantify specific sources of systematic uncertainties based on fixed input data provided to each participant. The blind predictions of crack growth life for the AFGROW Workshop were in very good agreement with the test data, and the predictions for the ERSI effort were generally within the statistical variation of the test data. However, the predictions of crack shape evolution did not show good agreement with the test results for either effort. The crack growth predictions for each crack direction were made using a single crack growth rate model based on the test specimen grain orientation and crack growth in the radial direction from the hole (L-T orientation). On further investigation, it was determined that different crack growth rate models (L-T and L-S) were required to predict crack shape changes as the initial corner cracks grew through the thickness of the test specimens. This paper will summarize the results of each blind round robin effort and compare the crack shape predictions made using single and dual crack growth rate models.

Alexander V. Litvinov, James A. Harter, Robert Pilarczyk

The Influence of Low and High-Cycle Fatigue on Dislocations Density and Residual Stresses in Inconel 718

The fatigue phenomena, as being one of the crucial importance for aircraft exploitation life, was investigated with different methods. The most important topic is to determine the state of the material and subsequently the moment of fatigue crack initiation. Many authors investigated the fatigue induced material damage and evolution of structure with different methods i.e.: scanning electron microscopy (Zhang et al. 2008), electron transmission microscopy (Hirsch and Whelan 1960), optical microscopy and two-beam interferometry (Kim and Laird 1978) or /and with modelling (Dingli et al. 2000). Nowadays, as the possibilities of diffraction methods were developed, so the conjoining of the diffraction image with the changes in the material structure can be applied to investigate the fatigue process in materials (Tahara et al. 2009; Huang et al. 2010).The aim of this work is to find out the relationship between the residual stress and dislocation density evolution applying the X-ray diffraction methods. The dislocations multiply and reorganize during monotonic and cyclic deformation, so their evolution can be a valuable information for investigation of fatigue phenomenon (Huang et al. 2008).The diffraction methods are non-destructive methods for quantitative analysis of grain level deformation (Korsunsky et al. 2004). In this study X-ray diffraction is employed to acquire the information about the evolution of elastic lattice strains and changes in dislocation density after fatigue cycling of Inconel 718 alloy.X-ray diffraction has been employed to assess the damage level under high and low cycle fatigue conditions and under the tensile test. The objectives of the work were achieved by two X-ray diffraction techniques: the analysis of residual stresses changes and investigations of changes of full width at half maximum of diffraction peaks which can be a measure of dislocation density changes. The diffraction results were compared to the hardness measurements.

Elżbieta Gadalińska, Maciej Malicki, Bartosz Madejski, Grzegorz Socha

Effect of Crack Length and Reference Stress on Variable Amplitude Fatigue Crack Growth Rate

The effective block approach is a concept where the crack growth rates are quantified and modelled by treating a block of variable amplitude (VA) loading as being similar to a single cycle of constant amplitude (CA) loading, but with a greater amount of energy compared to a single CA cycle. The fatigue crack growth rate (FCGR) per block can be determined for a certain VA spectrum and for lead cracks the crack growth rate shows a linear relationship with the crack length (exponential crack growth) and a cubed relationship with the reference stress (i.e. maximum spectrum stress) (Molent and Jones 2016). The applicability of the cubic rule for standardized test specimens is investigated by varying the reference stress between middle tension specimens. A short VA sequence from a combat wing root bending moment spectrum is applied to the specimens and is constantly repeated during the test. The crack length versus cycles curve is fitted directly to obtain accurate crack growth rates by assuming power law behaviour at all crack lengths and the presence of pivot points where the slope in crack growth rate is able to change at specific crack lengths. The resulting changes in VA long crack growth rates are correlated to the changes in crack length for a given reference stress and to the changes in reference stress between specimens. Translation of the VA FCGRs in the exponential crack growth regime for different reference stresses indicates that the exponent of the reference stress is 2.12. This is in contrast with an exponent of two on the reference stress if VA loading is governed by the reference stress intensity factor only or an exponent of three if the lead crack approach is applicable.

E. Amsterdam

Weibull or Log-Normal Distribution to Characterize Fatigue Life Scatter – Which Is More Suitable?

This paper is a follow-up to the Plantema Memorial Lecture that was presented by the author at the ICAF 2017 Symposium at Nagoya, Japan (Brot 2017). It will be shown in this paper, that there may be very large differences between the Weibull and Log-Normal statistical distributions of fatigue test results. These large differences have been confirmed using special software, that is commercially available. Fatigue test data from several sources have been combined to result in an 86-specimen database of fatigue life results. Analysis of these test results, using the software indicated above, determined that the Weibull distribution gave much more suitable results than the Log-Normal distribution.

Abraham Brot

Fatigue Life Enhancement Methods and Repair Solutions


Bonded Repairs of Composite Panels Representative of Wing Structure

In a collaborative effort, the Federal Aviation Administration (FAA) and the Boeing Company are assessing bonded repair technology of composite panels representative of transport airplane wing structure through test and analysis using the FAA’s Aircraft Beam Structural Test fixture. Emphasis has been placed on investigating methods and tools used to conduct analysis and predict structural performance of bonded repairs and those used to monitor and evaluate repair quality over the life of the part. The initial baseline phase of the program verified analysis models and provided an initial reference point for inspection and monitoring systems used to detect and track damage formation. Recent second-phase efforts support bonded repair size limit (BRSL) studies and methods used to predict the limit load residual strength for a failed scarfed repair in solid composite laminates. In general, methods under development for BRSL residual strength predictions correlated well with test results.

John G. Bakuckas, Reewanshu Chadha, Paul Swindell, Michael Fleming, John Z. Lin, J. B. Ihn, Nihar Desai, Erick Espinar-Mick, Mark Freisthler

Comparison of Rivet Hole Expansion for Protruding Rivets; Universal and with Compensator

Fatigue is one of the main reasons for airframe failure and cracks initiate often near rivets. Riveting technology influence strongly fatigue of joint. Rivet shank deforms sheets by expanding a hole and, if the force is high, compressive stresses are generated. It is the same phenomena as in cold working which is used to increase a fatigue life of elements with holes. The hole expansion, defined as increase of a hole diameter divided by an initial hole diameter, characterizes the degree of cold working. It can be used to assess a quality of a joint or even to estimate its fatigue life.The paper presents the numerical analyses of the hole expansion during riveting for two types of rivets, the universal rivet (MS20470) and the brazier rivet with a compensator (Russian branch standard OST 1 34040-79). The compensator is a small protrusion on the rivet head which is pressed into it during installation which results in improved hole expansion and increased fatigue life.The analyzed joint consists of two 1.5 mm sheets (2024-T3 alloy) and a 4 mm rivet (2117-T4 alloy). FE simulations of quasi-static riveting were performed with axisymmetric models for various force levels.Results show that hole expansion under the driven head is on the similar level for both types. Under manufactured head, the compensator caused much higher expansion. At the mating surface, which is a critical area for fatigue, expansion is significantly higher also. Moreover, hole expansion is much more uniform along the thickness for this type.Presented analyses and available results of fatigue tests convince the author that the concept of rivet with a compensator has significant potential to improve fatigue properties of joints practically without increasing costs.

Wojciech Wronicz

Effect of Alternative Paint Stripping Processes on the Fatigue Performance of Aluminium Alloys - Atmospheric Plasma De-painting

Typically, aircraft paint schemes lose their effectiveness for corrosion protection as well as cosmetic appearance every three to five years. As such, aircraft will undergo numerous removal and re-application cycles during their service lifetime to restore appearance, corrosion protection, or to enable inspection for fatigue cracks and corrosion damages. Current approved paint removal processes include chemical and abrasive media blasting. These processes yield high amounts of volatile organic compounds and generate large quantities of waste, which require proper disposal/treatment. They also have the potential to mask surface cracks and decrease the effectiveness of Liquid Penetrant Inspections (LPI). Concern over environment, safety and worker health with current paint removal processes has resulted in the enactment of new alternative removal processes during the past several years.Atmospheric Plasma (AP) has the potential to replace conventional paint stripping methods used for military aircraft structures in the Canadian Forces. As part of a Department of National Defence green initiative for aircraft repair and maintenance, NRC had been tasked to investigate the potential of this novel technology.In order for AP paint stripping to be accepted as an aerospace industry standard paint removal process, it must be thoroughly tested to demonstrate that it does not adversely affect the fatigue properties of the substrate. This paper investigates the effect of the AP paint removal process on fatigue crack nucleation and growth in aerospace aluminium panels.

Ali Merati, Marko Yanishevsky, Yan Bombardier

Effect of Strengthened Hole on the Fatigue Life of 2024-T3 Aluminum Alloy

Split-sleeve cold expansion processing was employed on the 2024-T3 aluminum alloy plate. Fatigue lives were compared according different expansion, and then the relationship of fatigue life and expansion was analyzed. Residual stresses were measured with different expansion, and the fatigue fractograph was analyzed by SEM. The results show that the split-sleeve cold expansion can obtain longer life compared with the non strengthened hole. The maximum fatigue life increased to 12 times with 6% expansion. When over 6% expansion, fatigue life began to decrease. The split-sleeve cold expansion can form beneficial residual compressive stress, and deferred the fatigue crack initiation. The fatigue fractograph shows quasi-cleavage fracture.

Hong Huang, Qingyun Zhao, Fenglei Liu

Fatigue Crack Growth in Pin Loaded Cold-Worked Holes

An experimental program has been carried out for the evaluation of the influence of the split sleeve expansion process on the fatigue crack growth in 2024-T351 aluminium alloy specimens. In previous ICAF Symposia, some information has been already given about the first part of the experimental activity (Helsinki 2015) and about the development of a numerical analysis method (Nagoya 2017). Both such preliminary papers referred to open hole specimens, tested under Constant Amplitude loading, while recently a pin loaded hole configuration has been evaluated. The non-inspectability of the configuration required some particular experimental effort, in order to collect test results in terms of crack growth as a function of number of cycles. To this end, a marker load technique was adopted: a block of high R ratio cycles (R = 0.9) was inserted in the R = 0.1 sequence, with the aim of obtaining information about the crack front evolution at different number of cycles.Due to the three-dimensionality of the residual stress field, which is of lower intensity at the face of mandrel entrance, a 1 mm radius quarter-circular notch was inserted by means of EDM on the mandrel entry side face. Moreover, it was necessary to include also mechanically milled notches, of similar dimensions.The results show a rather regular front evolution, and have provided important material for the development of accurate numerical methods, based on the evaluation of the residual stress field and on the subsequent modification of the stress intensity factor distribution along the corner crack front.The numerical analysis methodology is a specialization of the technique, already presented in the Nagoya Symposium, to the problem of single or double corner crack in a pin loaded hole. In particular, the strong three-dimensionality of the stress field poses a challenge to the block-by-block propagation analysis.

Luisa Boni, Daniele Fanteria, Domenico Furfari, Luigi Lazzeri

Fatigue Crack Propagation Influenced by Laser Shock Peening Introduced Residual Stress Fields in Aluminium Specimens

Laser Shock Peening (LSP) enables the generation and modification of residual stresses deep below the surface of metallic components. LSP-induced residual stress profiles provide penetration depths of compressive residual stresses in mm range, which can be used to retard the fatigue crack propagation (FCP) within thin sheets. These compressive residual stresses may lead to crack closure at significant applied tensile loads. This crack closure phenomenon is assumed to be one of the dominant mechanisms to reduce the load range at the crack tip, resulting in a fatigue crack retardation. This work provides an experimental and numerical investigation of the FCP in AA6056 based on C(T)100 specimens. Residual stresses were introduced by two-sided LSP treatment of the sheet material. The resulting residual stresses were determined by the incremental hole drilling method with electronic speckle pattern interferometry. The residual stress measurements on both sides of the specimens reveal differences of the residual stresses due to the laser shock peening process design. The occurrence of crack closure was evaluated by crack opening displacement vs. load curves, which can be used to determine the crack opening force. A multi-step simulation is applied to predict the residual stress field, the stress intensity factor range and rate if residual and applied stresses are present simultaneously as well as the FCP rate. Numerical predictions and measurements of the FCP rates are in excellent agreement.

Sören Keller, Manfred Horstmann, Nikolai Kashaev, Benjamin Klusemann

Influence of Bonded Crack Retarders on Damage Tolerance Performance of Fuselage Panel

The effect of bonded crack retarders on fatigue crack growth behavior and residual strength of fuselage panels was studied through the tests of 3 baseline panels (JZ-1, JZ-2 and JZ-3) and 3 reinforced panels (SR-1, SR-2 and SR-3). The baseline panels were curved fuselage panels with 7 stringers and 5 frames. As for the reinforced panels, crack retarders made of Glare-2 2/1 0.3 were bonded to the skin under each stringer and between adjacent stringers, along the direction of stringers. An initial circumferential skin crack with a length of 25.4 mm was introduced in the middle of each panel. Fatigue crack growth tests were conducted at axial constant amplitude loads till the crack tips approached the adjacent stringers. Static tests were performed as well to determine the residual strength of the panels with two-bay skin crack. Significant retardation was found for the reinforced panels with bonded crack retarders. With the same skin stress, the average fatigue crack growth life of reinforced panels was about 2.7 times of that of baseline panels. The residual strength of reinforced panels was over 37% higher than that of baseline panels.The fatigue crack growth behavior was also predicted based on finite element model and virtual crack closure technique. The residual strength was analyzed based on modified net-section criterion. Probable reasons for certain discrepancies were discussed.

Haiying Zhang, Dengke Dong, Yulong Wei, Weifeng Zang, Wenwei Yan

Is the Civil Aerospace Industry Ready to Implement Laser Shock Peening into Maintenance Environment? Questions to Be Answered and Minimum Requirements from Aircraft Manufacturer’s Perspective

Laser Shock Peening (LSP) system, involving complex set up and tooling, is not practical at all for in-service use (not compatible with airline maintenance constrains). To make LSP applicable at Maintenance Repair and Operations (MRO) and ensure reasonably simple setup and easy transportability to all around the world requires developing a “portable” device (i.e. low energy laser). The application of LSP as retrofit solution for in service commercial aircraft is particular challenging and currently no applications are reported. Applying LSP as a structural modification in critical component of in service commercial aircraft implies treatment at the MRO all around the world during already scheduled maintenance to avoid Aircraft on Ground situation, which can cost tens of thousands dollars a day. It is a common understanding that the depth of compressive residual stress over 1 mm can be achieved only if high energy laser (i.e. large laser spot) is used. It is demonstrated in this paper that low energy LSP system (≤ 200 mJ, pulse width of ≤ 25 nsec) and associated small laser spot size (< 1 mm diameter) can determine high compressive stress in the near surface of typical aeronautical Al alloy and compression depth above 1 mm. This residual stress profile is sufficient to extend the fatigue lives of critical components opening the door for development of portable LSP devices requiring low energy laser. The paper includes the investigation of low energy LSP system from residual stress characterization to fatigue life response of 7175-T7531 aluminum alloy. Finally, the authors review the minimum requirements of LSP portable device to ensure the compatibility with the operational environments typical of MRO.

D. Furfari, U. C. Heckenberger, V. Holzinger, E. Hombergsmeier, J. Vignot, N. Ohrloff

Fatigue Life Prediction at Cold Expanded Fastener Holes with ForceMate Bushings

ForceMate high interference fit expanded bushings, made by Fatigue Technology Inc. (FTI), are used by aircraft designers and maintainers to improve the fatigue and wear resistance of holes. While the fatigue life improvement (LIF) resulting from the installation of ForceMate bushings has been demonstrated experimentally, no analytical methods have been officially approved yet to take full benefit from the beneficial state of interference and residual stress resulting from this technology. To address this gap, a methodology was developed to analytically determine the LIF resulting from the installation of ForceMate bushings by explicitly taking into account the residual stresses and the effect of high interference fit bushings. To achieve this, a three-dimensional residual stress field is obtained from finite element process modelling of the ForceMate installation; fatigue crack nucleation lives are calculated; crack propagation analyses are conducted to calculate the resulting crack shapes and stress intensity factors, and the crack growth predictions are performed. This methodology was demonstrated on a CF–188 bulkhead at the holes attaching the main landing gear uplock mechanism. Based on this analytical study, a LIF of 6.6 was predicted for crack nucleation and 4.9 for crack growth from a 0.254 mm quarter-circular crack to a through-the thickness crack. While there are several aspects of this analytical study that need to be validated experimentally, the calculated LIF correlates well with the LIF typically obtained with the ForceMate system.

Yan Bombardier, Gang Li, Guillaume Renaud

Why Should We Encourage Usage of Interference-Fit Fasteners at Airframe Structural Joints?

It is well established that using interference-fit fasteners will obtain longer fatigue lives to airframe structures, compare to using transition-fit fasteners (or close-tolerance) and certainly clearance-fit fasteners. But, common practical manufacturing considerations, drive to less usage of the interference-fit fasteners (due to various installation difficulties of these fasteners being applied into the corresponding holes in the structure layers). In addition, concerns may be raised whether the fatigue advantage is actually being kept for any practical interference-fit installation method (or even fatigue disadvantage may occur due to installation procedures). It seems that there is lack of information regarding the influence on fatigue lives for the different practical installation methods of the interference-fit fasteners. This study presents testing results supported by analyses, for the influence on fatigue life, of the following two main parameters: (I) The fastener-to-hole fit level. (II) Two different common manufacturing practice for fastener installation methods of: hand plastic hammering and pneumatic steel hammering. The study shows that the fatigue advantage of interference-fit fasteners, is being kept even for the more aggressive installation method. The study results show that whenever fatigue life improvements are needed for structural joints, usage of interference-fit fasteners for these joints, is a good option to achieve it.

Carmel Matias, Ekaterina Katsav

Full Scale Fatigue Testing of Aircraft and Aircraft Components


Analysis Prediction and Correlation of Fiber Metal Laminate Crack Growth in Semi-Wing Full-Scale Test

This paper aims to demonstrate the correlation between simulation and experimental results obtained for artificially inserted crack propagation in the Full-Scale fatigue test of a semi-wing, developed by Embraer. The cracks were inserted on the lower wing skin, which was manufactured in Fiber Metal Laminates (FML), in order to validate the analysis methodology in this material. The application of cracks in a Full-scale test allowed the evaluation of several damage scenarios that could hardly be reproduced with high fidelity in panel tests, by incorporating: design details without simplifications and load redistribution among structural components. The analysis and test results will be substantiated and discussed.

Willy R. P. Mendonça, Danielle F. N. R. da Silva

Bombardier Global 7500 Fatigue Test Cycle Rate Commissioning to ¼ Life

This paper describes the commissioning campaign on the Bombardier Global 7500 durability and damage tolerance test (DADTT), as conducted by the experimental department of Bombardier Aerospace (BAEX). A rapid test rig commissioning time and smooth increase of cycle rate was a key objective in order to meet the ¼ life milestone for certification. Due to schedule changes, the time available for commissioning the loading apparatus and achieving the initial target ¼ life milestone was reduced from the original plan, with approximately three months’ time allocated to this goal. In anticipation of this aggressive testing milestone, the BAEX test team employed several new approaches and techniques to the DADTT that had not been applied on previous BAEX fatigue tests. A technical liaison from the National Research Council Canada was also on-site prior to and during commissioning activities.This paper opens with a summary of the objectives of the G7500 DADTT test conducted at BAEX, and focuses in particular on the techniques and approaches used by the BAEX team to achieve a rapid, efficient and productive test commissioning phase, such that the target cycling rate was achieved in a faster timeframe than anticipated. These techniques included: judicious data-informed hydraulic and pneumatic hardware selections; informed design choices to minimize mass and actuator count; hydraulic and load controller training and procedure generation on a dedicated independent test platform; extensive hardware-in-the-loop tuning to maximize performance; and using the global finite element model (GFEM) of the test article, coupled with a simple pneumatic model, to better estimate initial test load transition times.As a result of the techniques employed, the ¼ life milestone was successfully achieved, contributing to the final certification package for the world’s largest purpose built business aircraft, which received Type Certification from Transport Canada on September 28, 2018.

C. André Beltempo, Alexandre Beaudoin, Robert Pothier

Changing the Philosophy of Full-Scale-Fatigue-Tests Derived from 50 Years of IABG Experience Towards a Virtual Environment

IABG has continuously developed new techniques in order to improve time, cost and quality for full-scale fatigue tests.With the advancement of virtualisation of the aircraft development and certification process, questions have to be raised as to how full-scale fatigue tests can be incorporated into these processes and how virtualisation can benefit from physical testing.Virtualisation will speed up aircraft development to its next level and will change time, cost and planning expectations. Will full-scale fatigue testing (FSFT) still fit into this context?Authorities are demanding physical tests for new aircraft types for good reasons. Testing experience in fact confirms that some significant issues were first detected during the full-scale fatigue tests.The purpose of full-scale fatigue tests may, however, be changing. The amount of available but unused information gathered during fatigue runs is still sizable, and only by changing the approach to fatigue tests the return on investment could be increased.It will be required to introduce agile methods and processes to the preparation and performance of the full-scale fatigue test to align the test with the dynamic evolution of the aircraft’s requirements and design. In the end this could result in a new philosophy of full-scale fatigue tests, moving from flight-by-flight testing towards purely artificially triggered load sequences in order to demonstrate the correctness and reliability of the virtual qualification tools which will be of crucial importance for any virtually based certification. This paper, however, is aiming as well at describing some of the more near-term benefits the FSFT can contribute in view of the long-term goal.

Gerhard Hilfer, Olaf Tusch, Don Wu, Michael Stodt

Combined Static and Fatigue Tests of the Full-Scale Structure of a Transport Aircraft

The article discusses a non-traditional approach in strength tests of the full-scale structure of a transport aircraft, which consists of combining static and fatigue tests on one object. The test object included a full span wing with installed pylons, the middle part of the fuselage and the main landing gear. The tests were carried out on the bench, which allowed reproducing both static cases of loading and variable loads of flight cycles. To confirm the static strength, the structure of wing half span was loaded with limit loads and simultaneous strain measurement. The data of strain gauges verified the finite element model (FEM), on the basis of which the prediction of the stress-strain state of the structure was made with the ultimate loads. To confirm the strength capacity of the upper wing panels under the stability conditions, tests of full-size wing panels were carried out. After loading with the limit load, fatigue tests were carried out which are required to confirm the service life.

K. S. Shcherban, A. A. Surnachev, M. V. Limonin, A. G. Kalish, O. V. Chuvilin

Conception of Modular Test Stand for Fatigue Testing of Aeronautical Structures

Fatigue tests of specimens and components are a necessary part of structural development in aerospace but they are expensive. It can be a problem especially in the case of low cost projects or students researches. Most of them are conducted on testing machines with simple specimens, usually with loads limited to tension mode.The paper presents the Modular Test Stand which was design and developed to decrease the cost of fatigue tests and to test specimens with more complex load condition. The stand consists of three identical sections which are structures similar to the airframe, namely the wing box. Sections are connected, and during a test are loaded in the same manner by bending or twisting moment. The whole section structural node, a particular joint or a skin can be an object of testing.Based on FE calculations, the design of the stand was developed. The desired requirement were uniform stress distribution in skin panels and axial stress level during bending equal to 100–120 MPa. Two stands were constructed - one for bending and one for torsion. Displacements and shearing strains were measured in the central part of the middle skin panel during torsion with the use of Digital Image Correlation method. The measurement correlated very well with FE calculations and confirmed uniform strain distribution in the panel.The stand can be used to examine joining methods, materials but also structures with damages or repairs as well as various types of SHM sensors. The main advantage is a possibility of testing up to six specimens at the same time (double side of three sections) which reduces the cost of a single test. Additionally, a more complex load state can be achieved compare to simple specimens.

Andrzej Leski, Wojciech Wronicz, Piotr Kowalczyk, Michał Szmidt

Full Scale Fatigue Testing for Mitsubishi Regional Jet

Mitsubishi Aircraft Corporation is performing the full-scale fatigue testing (FSFT) for Mitsubishi Reginal Jet (MRJ) type certification. Main objective of this test is to show freedom from wide spread fatigue damage (WFD) during the life of aircraft and establish Limit of Validity (LOV). Prior to the test, WFD susceptible structures are defined based on the stress distributions and structure configurations, and they are fully covered by this test. Test duration of FSFT is 240,000 flights (3 × DSG of 80,000 flights). The flight-by-flight loading spectrum is newly designed for MRJ and its loads occurrence data was verified by flight test data. Furthermore, to reduce the test duration, low loads omission was applied based on the results of some spectrum verification tests. During fatigue test, scheduled inspection consistent with MRJ maintenance program is planned.Since the main objective of FSFT is no WFD substantiation, no artificial crack is introduced during FSFT. Therefore, damage tolerance evaluation (i.e. crack growth analysis validation) is separately conducted by sub-component level testing. For example, damage tolerance substantiation for fuselage structure was performed by using curved panel test facility. This facility can simulate the pressurization load with axial load and their loading sequence can be customized for each test. Major detail design points of fuselage structure such as the lap/butt-joint, cut-out structure and repaired structure are individually evaluated by this type of tests. Based on the test results, potential fatigue critical locations and crack growth behaviors are efficiently investigated, and they significantly contribute to the crack growth analysis validation necessary for CFR/CS 25.571 compliance.

Koji Setta, Toshiyasu Fukuoka, Kasumi Nagao, Keisuke Kumagai

Full-Scale Fatigue and Residual Strength Tests of the Composite Wing Box of a Passenger Aircraft

Fatigue and residual strength tests were performed of the wing box, which was manufactured using a polymer composite material according to a new infusion technology. There are investigated influence of the technological effects (porosity, delamination, etc.) and operational (impact) defects on the durability and strength of the composite structure.

K. S. Scherban, A. Yu. Zakharenkova, V. V. Konovalov, S. V. Kulikov, V. E. Strizhius

Full-Scale Fatigue Testing from a Structural Analysis Perspective

Generally speaking, full-scale fatigue tests are used to demonstrate ‘Means of Compliance’ (MoC) for Type Certification. Aircraft are designed in accordance with fatigue and damage tolerance requirements; the main purpose of the fatigue test being to provide the physical evidence necessary to validate design assumptions.Located at the top of the test pyramid, these tests come with significant investment. The test objective is therefore not only limited to compliance with regulations, but also aims to obtain highly important experience of the airframe providing significant benefit to future applications.This paper presents the structural analysis view of full-scale fatigue testing, which drives large parts of the test definition, execution, exploitation and use of test outcomes. Evolution of the general fatigue test approach, alignment of fatigue requirements with test execution and exploitation of the test results for Airbus aircraft is explained. Finally, the paper also aims to capture details regarding future development of full-scale fatigue testing.

Derk Daverschot, Paul Mattheij, Mathias Renner, Yudi Ardianto, Manuel De Araujo, Kyle Graham

Hawk Mk 51/51A/66 Tailplane Full-Scale Fatigue Tests

Until 2017, there was no certainty about the fatigue life of Hawk tailplanes in FINAF’s flight conditions. Then full-scale fatigue tests were performed to determine if the FINAF is required to procure more tailplanes, and to extract evidence, which could be used to increase the structural inspection interval times. The tests were executed with two 4000 FH flown tailplanes and the goal was to achieve additional 2000 FH with a scatter factor of 5. Test loads were applied with actuators feeding both buffeting and maneuvering symmetrically at the same time. Test’s spectrum was based on the FINAF OLM strains and on the usage spectrum of the FINAF flights 2014–2015.Limited NDIs were done after every 200–340 EFH and full inspections after every 1000 EFH. Several damages, such as broken rivets and cracks in spars and angles, arose. Following the testing, the tailplanes were subjected to RSTs with the load corresponding the ultimate design load. The tailplanes passed the RSTs without noticeable additional damages. Centre sections were torn down for more detailed inspections. Some fault indications were obtained from the buttstraps, but all the defects were very small. Seven cracks were found on the skins and one location could be determined as the critical location.The centre joint survived the test period. The residual strength was sufficient with a 20 mm crack at the skin rivet hole, which was estimated to be the most loaded. The tests gave solid basis for increasing the TP’s acceptable usage life by 1000 FH. It was possible to determine the crack propagation rate to verify the structural inspection period to be applied. Considerable cost savings will be achieved, because the inspections can now be optimized. In addition, now it is known that the current number of TPs is sufficient with the additional 1000 FH for the targeted HW life cycle, and no additional procurement is required.

Risto Laakso, Jussi Kettunen, Juha Lähteenmäki

Progress on the Pathway to a Virtual Fatigue Test

The Advancing Structural Simulation to drive Innovative Sustainment Technologies (ASSIST) collaborative program was initiated by Defence Science and Technology (DST) Group to foster improvements in fatigue life prediction technologies, with the long-term goal of a virtual fatigue test. It is based upon a growing series of airframe challenges, where participants are invited to test state-of-the-art fatigue prediction technologies on problems based on real aircraft structures and loads. Since each challenge is underpinned by a set of demonstrated test results, the predictive ability of all technologies can be accurately assessed. Furthermore, the collaborative forensic review of the challenge results is considered a key output, which will allow the limitations of predictive techniques to be better understood and addressed in future research.DST has established a collaborative online space for the ASSIST community at . It is being used to share essential data associated with each challenge, provide a collaborative space for discussions, and post results and evaluations. The first ASSIST challenge has been completed and two further challenges have been released to the ASSIST community. The first ASSIST challenge, which was based on a fighter wing-root shear tie post, is described here. It demonstrated the critical importance of having accurate fatigue crack growth data, in addition to understanding the fatigue crack shape and local stress field when the crack depth is still very small. It is considered that future ASSIST challenges will provide many further insights, which will drive improvements to current fatigue life prediction technologies. It is envisaged that the growing database of ASSIST airframe challenges will provide an understanding of the accuracy that can be achieved with current fatigue life prediction methodologies. This is another important product of the program, because it enables the adoption of such methodologies on real aircraft structures.

Ben Dixon, Madeleine Burchill, Ben Main, Thierry Stehlin, Raphaël Rigoli

Testing Approach for Over Wing Doors Using Curved Fuselage Panel Testing Technology

The A321neo ACF (Airbus Cabin Flex configuration) contains newly designed Overwing Doors (OWD) that provide an automatic opening function for the case of evacuation. Due to the significant structural differences between the previous “hatch” design used on A319 and A320 aircraft and the new OWD design it has been decided to test the OWD’s and the surrounding fuselage structure to demonstrate the Fatigue and Damage Tolerance capabilities of the structure.This demonstration should be done by means of a fatigue and damage tolerance test. An appropriate test set-up had to be selected. Two generally different approached were investigated. This selection process led to the decision to follow the curved fuselage panel test method.Curved fuselage panel testing was developed in order to test undisturbed, regular panels. During the last years, this method was improved to allow testing of panels with major non-regularities, for instance door cut outs, floor beams or similar features.

Mirko Sachse, Matthias Götze, Silvio Nebel, Sven Berssin, Christian Göpel

Very High-Cycle Fatigue Characteristics of Cross-Ply CFRP Laminates in Transverse Crack Initiation

Fan blades are subjected to very high-cycle loadings during the design life, so it is essential to evaluate the giga-cycle fatigue characteristics of carbon fiber reinforced plastic (CFRP) laminates. In this study, the transverse crack initiation of the cross-ply CFRP laminates in very high-cycle fatigue region was evaluated using an ultrasonic fatigue testing machine. The fatigue tests were conducted at the frequency of f = 20 kHz and the stress ratio of R = −1. In order to suppress temperature rise of the specimen, the intermittent operation with the loading time of 200 ms and the dwelling time of 2000 ms was adopted. The fatigue life data to transverse crack initiation in very high-cycle fatigue region was compared with the data of the fatigue test which was conducted at the frequency of f = 5 Hz and the stress ratios of R = 0.1 and −1 using a hydraulic control fatigue test machine. It was evaluated considering the influences of the stress ratio and the thermal residual stress by using the modified Walker model. The fatigue life to the transverse crack initiation of the cross-ply CFRP laminates in the very high-cycle region exceeding 108 cycles was on the extension of the test data in the low cycle region.

Atsushi Hosoi, Takuro Suzuki, Kensuke Kosugi, Takeru Atsumi, Yoshinobu Shimamura, Terumasa Tsuda, Hiroyuki Kawada

Application of Optical Fiber-Based Strain Sensing for the Full-Scale Static and Fatigue Tests of Aircraft Structure

An optical fiber-based Rayleigh backscattering distributed strain sensing system was adopted as the main structural integrity monitoring tool for airframe Full-Scale fatigue and ultimate tests. The strain signature along all major structural elements, as measured by the optical fibers, at each loading step was recorded and analyzed in real time. A specially developed human interface enabled easy tracking of emerging damage-related non-linear phenomena. This sensing concept reduces the need for adding hundreds of electrical strain gauges and eliminates intermediate conventional structural inspections during test, all leading to reducing test duration and cost.

U. Ben-Simon, S. Shoham, R. Davidi, N. Goldstein, I. Kressel, M. Tur

In-Service Experience, Life Extension and Management of Aging Fleets


Analysis of Adhesive Disbond Occurrences in Rotor Blades of Mi-2 Helicopters

Polish Air Force Institute of Technology (AFIT) is responsible for delivery service life extension program for the main rotor blades of multiple helicopters in the Polish Air Force (such as Mi-2, Mi-8/17, Mi-14, Mi-24). Due to the operational profile, certain number of blades got decommissioned when the calendar service life is reached, while the accumulated flight hour count is still below the hourly service limit. This is why a dedicated research program, as well as inquiry into the usage and technical condition of the blades is necessary. To maintain such a program, regular inspections of the blade structure are required.In the paper, the Mi-2 (hoplite) service life extension program findings are presented. The Mi-2 rotor blade service life extension program has been implemented in three phases, and life extension of 42 months was achieved. The program is based on detailed non-destructive tests. Inspection procedures include verification of the spar structure (corrosion, cracks) as well as the verification of adhesive bonded joins between various elements of the blade (skin and spar, skin and core, etc.). The inspection is performed with the use of automated NDE systems with C-Scan data presentation possibility. Because adhesive disbonds are the main reason for component rejection and decommissioning, they are the main focus of the paper. A study of more than hundred non-destructive inspections is included in the paper. The inspection methodology is presented, as well as the statistics and distribution of the blade failure modes. The adhesive disbonds are also characterized. The study aims to determine the factors that drive the disbond process, and environmental factors are taken into account.

Piotr Synaszko, Krzysztof Dragan, Michał Sałaciński, Mirosław Wrona

Approach to Evaluation of Delamination on the MiG-29’s Vertical Stabilizers Composite Skin

The airframe of the MiG-29 jet fighter is a semi-monocoque, aluminum structure. The vertical stabilizer skin is the only safety-critical load bearing components made of composite materials. According to the manufacturer’s instruction, the vertical stabilizer operation is based on the safe-life approach i.e. monitoring of hourly and calendar based service lifes. In addition, visual inspections including tap-testing are periodically performed.The service life of the MiG-29s operated by the PAF has already been extended: calendar life – from 25 to 40 years and hourly life – from 2000 to 4000 flight hours. In the course of the extension program, additional maintenance and overhaul actions had to be introduced.The additional actions consisted of structural monitoring of the stabilizer airframe – especially skin delamination and stringer debonding. Non-destructive testing has been used for this purpose (especially ultrasonic methods). Damage tolerance for these defects (delamination and debonding) was based on occurrence statistic (distribution) obtained from the aircraft in service.Based on the research from the last decade, it has been established that the skin defects do not propagate, however the number of defects increases. Recently, defects exceeding the statistic-based criteria were also observed. Therefore a new approach for damage criteria has been implemented, based on Finite Element simulations of the defect initiation and evolution – for which fracture mechanics methods were used.This article contains information on damage modeling and methodology for checking the condition of damaged structure.

Michał Sałaciński, Piotr Synaszko, Dawid Olesiński, Piotr Samoraj

Evaluation of a PC-9/A Wing Main Spar with Misdrills Using Enhanced Teardown at Resonance

Widespread production misdrills in Royal Australian Air Force (RAAF) PC-9/A wing main spar lower caps were a potential threat to airworthiness and fleet availability late in the service life of the aircraft. A rapid, novel, enhanced teardown of a retired RAAF PC-9/A wing spar with production misdrills was successfully completed to address this risk. Dynamic block loading was applied to the main spar lower cap of a wing with multiple misdrill indications identified in-service using x-ray inspection. Using an enhanced teardown experimental method and cycling the wing at resonance, fatigue failure occurred within the test section from a main spar misdrill in less than fifty cumulative hours of cycling. A damage tolerance prediction, testing results, teardown inspection and forensic assessment of the spar lower cap was combined to deliver timely and valuable risk mitigation advice for PC-9/A fleet structural integrity managers. Greater confidence in existing fleet management was fostered in a very short timeframe and the opportunity presented by a retired wing with misdrills was fully leveraged. In addition, an innovative structural experimentation method for aircraft sustainment was pioneered.

Ben Main, Keith Muller, Michael Konak, Michael Jones, Sudeep Sudhakar, Simon Barter

Holistic Approach for Determining a Helicopter’s Airframe Interval for Depot Induction

The US Coast Guard inducts MH-65 helicopters into their programmed depot maintenance (PDM) overhaul approximately every 48 to 52 months. The primary purpose of the PDM is to inspect for structural corrosion and fatigue, and if discovered, to repair the compromised structure. There are aircraft, however, that are prematurely—and thus unnecessarily—inducted into PDM because the aircraft are neither corroded nor are they cracked, resulting in a needless but expensive overhaul evolution. Other aircraft arrive in depot with corrosion beyond repair limits due to harsh and extreme environments of operations. The PDM interval is USCG mandated based upon the historic amount of corrosion observed, not from prescribed interval recommendations from either the Original Equipment Manufacture or FAA recommendations. A holistic evaluation and assessment of the MH-65 programmed depot maintenance process results in the following major findings and recommendations: A component criticality and prioritization process determined the potential structural depot drivers for PDM induction. For the PDM drivers, the components were evaluated for the following three damage progression failure modes: (1) fatigue, (2) pitting and fatigue, and (3) general corrosion degradation. Corrosion & Fatigue Damage Tolerance Assessments (CDTA) were performed to yield the depot induction interval capturing the full helicopter airframe structure. The results of this work recommends increasing the mean PDM induction interval to 60 months ±10% as the optimal depot interval. Significant cost avoidances are afforded by implementation of the PDM extension. The project consisted of a comprehensive study of the many aspects of the induction cycle, addressing the “who, why, what, where, and how” of the PDM. The primary focus of this paper is the qualitative and quantitative methods employed to determine the better “when” the induction interval should be to maintain the fleet airworthiness.

Craig L. Brooks, Samuel Benavides

Structural Health and Structural Loads Monitoring


A Guidance to Derive Statistical Data for Asymmetrical Maneuvers on Transport Operation

The airplane fatigue design requires statistical dataset representative of the loading relevant conditions that affect the structural sizing. The current work aims to propose a guidance to derive flight parameters statistical dataset to represent asymmetrical maneuvers on transport operation. The guidance comprises the definition of an acceptable sampling rate, the peaks and valleys counting method and the criteria for flight segmentation into flight phases. Moreover, such guidance indicates the better flight parameters to represent asymmetrical maneuvers, leading to a unique statistical dataset for airplanes with the same flight qualities construction concept.

Juliana Diniz Mattos, Diego Silva Peixoto, Frank Machado

A Machine Learning Approach to Load Tracking and Usage Monitoring for Legacy Fleets

With changes to aircraft usage due to expanded roles, operators need to monitor the usage of the aircraft and component loads to ensure safe operation of the components within their fatigue lives. Accurate load monitoring of aircraft component loads during flight is a challenge that has inspired the implementation of computational intelligence and machine learning techniques to replace the need for costly sensor systems. Much research has been carried out to develop a methodology for load signal and fatigue life estimation using only data from standard flight state and control system parameters. The National Research Council’s approach to load and usage monitoring is centered on leveraging the data recorded by existing instrumentation using machine learning models and data mining techniques. This flexible approach has been demonstrated on several helicopter platforms manufactured by different original equipment manufacturers. Results from the Australian S-70A-9 Black Hawk and Canadian Forces CH-146 Griffon (Bell 412) are provided to demonstrate the analysis and outputs enabled by accurate load monitoring. These outputs include load signal estimates, fatigue damage accumulation, and load exceedance plots and comparisons.For many operators, access to component material information that supported initial design analysis is very limited, so the provided outputs do not assume that this type of information is readily available. Using this analysis, valuable insights can still be obtained, particularly through load exceedance comparisons. Accurate load and usage monitoring is a universal goal that is not easily achieved on older aircraft without the sophisticated and powerful health and usage monitoring systems that are available today. This research targets these smaller legacy fleets to help satisfy their usage monitoring requirements and ensuring structural integrity.

Catherine Cheung, Srishti Sehgal, Julio J. Valdés

Structural Integrity Control Technology Based on Structural Damage Monitoring

Structural integrity is an important attribute of aircraft structure. Usually, the description of structural integrity is a qualitative form, which emphasizes on the keep up of structural integrity. But, the quantitative description of structural integrity is help for carry out more comprehensive work for keeping up structural integrity. In this study, the concept of structural integrity control and a quantitative measurement method that uses structural integrity degree to measure structure integrity are proposed. And the structural integrity degree is evaluated based on durability degree, availability degree, safety degree and livability degree. Then, an aircraft structural integrity control technology based on structural damage monitoring is introduced. The implementation method and characteristic advantages of aircraft structural integrity control technology based on structural damage monitoring are expounded. Finally, aiming at the characteristics of military aircraft, the concepts of structural battle integrity and recoverability are established, and its measurement and control methods are briefly described.

Yuting He, Teng Zhang, Binlin Ma, Xianghong Fan, Xu Du

Application of RLC Filters and Analog Circuits for Increasing Information Bandwidth of Channels of Data Acquisition Units

Fatigue cracks can be detected and monitored by using resistive crack propagation sensors (CPS). Usually the sensor is composed by a series of conductive strands connected in parallel. When the crack propagates underneath the sensor, subsequent strands of the sensor are breaking down which results in increase of the sensor’s resistance. Therefore, by tracking the resistance of the sensor, the condition of the structure being monitored can be inferred. However, due to nonlinear behavior of the parallel connection resistance, usually CPS sensors are unidirectional only, unless the geometry of individual sensor’s strands is designed properly. Also, the result of the measurement can be influenced by many factors, e.g. the temperature, acting not only on the sensor itself, but also on other elements of the measuring system, which can bias the results and increase risk of false calls. In the paper the design of an analog device capable to track integrity of every strand of a sensor or make measurements of several CPS using single analog input channel is presented. The underlying idea is based on properly designed RLC band-pass filter. An example of the implementation and computer simulations as well as results of first laboratory tests are also delivered in the paper.

Kamil Kowalczyk, Michal Dziendzikowski, Artur Kurnyta, Patryk Niedbala, Krzysztof Dragan

Effect of Plate Thickness and Paint on Lightning Strike Damage of Aluminum Alloy Sheet

Aircraft usually avoid areas associated with lightning. However, when such areas are directly along the takeoff or landing path and cannot be avoided without delay, there is a risk of lightning strike during flight. After an aircraft has been hit by lightning, maintenance technicians have to identify the lightning entry and exit points and conduct any necessary repairs before the next flight, and must perform these tasks as quickly as possible to minimize disruption to operations. The metallic material has high conductivity and it is known that the damage is small comparing to composite material. However, the knowledge concerning about the lightning strike damage on metallic material contributes to minimize the damage size and the duration of repair.The effect of plate thickness on lightning strike damage for 2024-T3 Aluminum alloy sheet is evaluated. Test result shows that the surface of the specimen is melt by the high heat input during the lightning strike and the area can be distinguished by the appearance. In addition, the result shows that the lightning strike damage on the specimen with paint becomes small comparing to that on the specimen without paint, and then the paint acts as insulator.

Takao Okada, Hiromitsu Miyaki, Yoshiyasu Hirano

Evaluating the Influence of SHM on Damage Tolerant Aircraft Structures Considering Fatigue

First patents concerning the integration of Structural Health Monitoring (SHM) capabilities into aircraft have been granted over 70 years ago (Green 1948). Since then, numerous potential applications of SHM along the entire aircraft lifecycle have been identified in scientific literature. However, the application of this technology in commercial aviation is currently limited to field studies carried out primarily within aging fleets. Aside from recent advances in sensor technology, growing research effort has been committed to identify and quantify promising business cases for SHM (Bos 2017). Even though numerous individual applications have been investigated with regard to their lifecycle cost impact, current studies yield varying results. In order to facilitate the integration of SHM in commercial aviation we plan to put forward an integrated analysis framework for SHM technologies covering the entire aircraft lifecycle including design, operation and retirement. As part of this framework the work at hand introduces an expanded evaluation approach based on (Schmidt and Schmidt‐Brandecker 2009) to identify the impact of SHM on aircraft structural design considering structural weight, risk of structural failure and inspection intervals. The proposed methodology is limited to damage tolerant structures prone to fatigue and considers only component sizing rather than the redesign of structures. By using the established relation between structural dimensions, weight, inspection intervals and aircraft end of life, the approach is calibrated with inspection intervals provided by the manufacturer. Subsequently, the influence of SHM can be analyzed using an ideal sensor system (no false alarms). Finally, a case study is presented demonstrating the suggested methodology using the example of a commercial narrow-body passenger aircraft.

Dominik M. Steinweg, Mirko Hornung

Fatigue Crack Growth Approach for Fleet Monitoring

Aircraft are subjected to unpredictable variable amplitude loadings (operational loads, varying widely according to the missions flown by the aircraft). For safety reasons, it is necessary to estimate each aircraft’s actual fatigue “consumption” relative to its potential. For this purpose, most French military aircraft are equipped with load monitoring systems, giving direct or indirect access to the in-service loads. The cumulative fatigue damage is calculated at various points of the structure pointed out as being critical. Thus, for fleet life-extension purposes and usage predictions, the use of more precise damage and crack growth prediction models is a major concern.The initiation models are yet insufficient to cover the entire aircraft lifetime. In order to maintain the design safety margin, the airworthiness authority introduced the damage tolerance philosophy: cracks are allowed below a given size provided that the propagation process is well known. Systems currently in operation on the French Air Force enable the acquisition of large number of flight parameters. This data associated with crack growth models provide information about the propagation phase. The fatigue potential consumption during the operational life can thus be estimated by combining initiation and propagation models.Initiation model are based either on real load sequence combined with unit damage matrix or on rebuilt load sequence (from g-counters output) combined with cumulative damage laws. Different propagation models are investigated. At this stage, experiments have been carried out on CT samples and are ongoing on a geometry close to the spar one (thick holed aluminum alloy part).Further investigations will be carried out to extend the conclusions made on aircraft equipped with load monitoring systems to aircrafts only equipped with g-counters. It means build a temporal spectrum using g-counters output in order to take load history into account in the propagation process (mostly retardation effect due to variable amplitude loading).

Olivier Gillet, Bastien Bayart

Flight Testing of an Ultrasonic Based SHM System

We have been developing an Ultrasonic based structural health monitoring (SHM) system which can provide essential information on structural integrity of airframes. The information can contribute to create novel design philosophies, improve inspection activities in manufacturing processes, and optimization of aircraft operation with the condition-based maintenance. In the SHM system, ultrasonic Lamb waves are measured and analyzed to evaluate damages in airframes because ultrasonic Lamb waves changes with the changes of structures such as damage initiations and their growth. In order to achieve implementation of the SHM system to actual operating aircraft, we have conducted various types of tests for over 15 years, in which we assessed damage detection capability, accuracy and probability of damage detection, environmental durability, and applicability to aircraft system, and so on.In this report, results of a flight testing as a pre-trial toward the implementation of the SHM system were summarized. In the flight testing, we evaluated influences of environmental conditions to measurements of ultrasonic Lamb waves and detection capability of the SHM system in actual aircraft operating conditions with a flying test bed owned by Japan Aerospace eXploration Agency (JAXA). From the results of the flight testing, it was confirmed that ultrasonic waves could be measured in flight conditions just like on ground and also damages introduced to the test specimen could be detected by analyzing the changes in waveforms of the ultrasonic Lamb waves even in the actual operating aircraft.

Hideki Soejima, Takuya Nakano, Makoto Yokozuka, Yoji Okabe, Nobuo Takeda, Noriyuki Sawai

Lightning Strike Damage of CF/Epoxy Composite Laminates with Conductive Polymer Layers

Carbon fiber reinforced plastics (CFRPs) are prone to severe damages by lightning strikes due to their low electrical conductivity. Current lightning strike protection (LSP) technology generally consists of metal foils/films on the surface of composite airframe structures. The present work aims to introduce all polymeric conductive layer for LSP of CFRP structures. Intrinsic conductive polymer i.e. Polyaniline (PANI) is used to make a thermosetting polymer mixture. CFRPs coated with this electrically conductive, all polymeric layer was tested against simulated lighting strike. It has been shown that PANI-LSP specimens dissipated the lightning current effectively and provided enough residual mechanical properties of CFRPs.

Tomohiro Yokozeki, Vipin Kumar, Yu Zhou, Takao Okada, Teruya Goto, Tatsuhiro Takahashi

Machine Learning Application on Aircraft Fatigue Stress Predictions

Machine learning applications are rapidly gaining momentum and already offer significant benefits in many areas of industry; from traffic predictions and flow of goods to fighting cancer, the world is turning increasingly to intelligent algorithms to overcome its most persistent challenges. With an ever-growing demand for air travel and an increasingly competitive market, it has now become a priority for commercial aircraft manufacturers to join the digital revolution.The fundamental requirement for tailored aircraft maintenance solutions providing fully optimised scheduling without compromise to safety is knowledge of real-world aircraft usage. By harvesting fleet-wide aircraft usage parameters (e.g. ‘big data’ flight-by-flight recordings, meteorological conditions, etc.), and exploiting them with the application of validated machine learning algorithms, highly accurate predictions of the internal loading conditions of a structure become possible based on measured and recorded aircraft parameters (e.g. speed, altitude, flight-configurations, accelerations etc.). The major benefit of such data analytics is the possibility to recreate the real loading sequence as experienced by the structure. To fully validate any conclusions from machine learning, it must be recognised that other information will need to be considered in the context of the analysis, e.g. additional in-service data, full scale test results, theoretical analyses, etc.The significance of the potential capabilities held by machine learning applications to predict aircraft structure internal load distribution is a major enabler for linking and comparison of practical aircraft usage against that of any average fleet or initial assumptions; the unlocking of such potential offers game-changing benefits in civil aviation.The main input needed to make relevant and reliable predictions is a comprehensive dataset. Generally speaking, this and most other parts of the puzzle are already available; the main challenge for machine learning will be to integrate and validate these parts harmoniously, obtaining new predictive capabilities by enhancing analytical power.

Eugene O’Higgins, Kyle Graham, Derk Daverschot, Julien Baris

Modernizing the A-10 Loading Spectrum Development Process

As the A-10 fleet transitioned away from the legacy MXU recorders that were becoming increasingly obsolete and toward a new recording solution, the downstream tools used to process the new data also needed replacement. The initial focus for these software tools was to increase their flexibility and capability over the legacy programs in understanding the details of how the A-10 fleet was being flown. This focus allowed for many valuable studies regarding issues such as relative severity across the fleet, gunfire rates, stores carried, etc., and how that impacted the structural integrity of the fleet. After the new recorders had been flying on a subset of the fleet for a few years, the decision was made to install them on all remaining A-10s. This was to not only benefit the quality and quantity of data going into the Loads/Environment Spectra Survey, but also to improve the Individual Aircraft Tracking Program, which by this point was dealing with obsolescence issues of its own. The order of magnitude increase in data to be processed led to the data processing and maneuver spectrum development functions being automated and transitioned to the Tinker AFB ASIMIS office, giving the USAF even more flexibility and internal capability to gather usage data necessary for force management. This paper gives the history of this transition from legacy recorders and tools to modernized processes and organic capabilities, and the program benefits that have resulted from that transition.

Luciano Smith, Mark Thomsen, Devin Butts, Kurt Schrader

Nondestructive Evaluation for Damage Tolerance Life Management of Composite Structures

Polymer matrix composites (PMCs) are experiencing a growth in their use for civilian and military aircraft. However, the certification process for PMCs leads to new requirements for nondestructive evaluation/inspection (NDE/I). For example, in metallic structures current practice as defined in MIL STD 1530Dc1 uses slow crack growth analysis requiring the NDE/I technique to have a defined probability of detection (POD) curve to enable risk calculations. However, certification processes used to date for PMCs are closer to safe-life methods. However, there is a desire to alter the approach for managing PMCs structures in the US Air Force (USAF) to follow slow damage evolution criteria as is done for metallic structures today. To realize this desired capability, predictive modeling is being developed for slow damage evolution in PMCs. As a key input to these models, metrics of damage are required from NDE/I-based methods that characterize the geometry of the defects. Explored technical approaches include conventional pulse-echo methods, including resolution of tip diffraction from delaminations at the individual ply level, but amplitudes of these responses are quite small. An alternative approach uses localized single-sided pitch-catch methods to evaluate ultrasound propagating through and around a damaged region, such as typical damage from impacts. It uses the signal transmitted through damaged regions to extract features indicating matrix cracking or internal geometric attributes of delaminations that supplement the conventional one-sided pulse echo measurement. The methods includes using model-based methods and feature extraction approaches. Initial results show significant promise and damage state verification is obtained from destructive methods based on serial sectioning. This provides ground truth for measurements and is a reference for metrics of performance. With this capability, the Building blocks are in place to realize slow damage growth damage tolerance for PMC structures.

Eric A. Lindgren, John C. Aldrin, David H. Mollenhauer, Mark D. Flores

Perspective of Structural Health Monitoring for Military Aviation in Poland

About 40 years have passed from introducing non-destructive testing inspections (NDI) as an inherent component of damage tolerant approach in order to ensure structural integrity of aircrafts. Over the years, non-destructive testing (NDT) methods became very accurate and reliable in damage detection and assessment, allowing to achieve very high level of safety in the aerospace. However, still there are some issues of this aircraft design paradigm which need to be addressed in the future. First, NDI are scheduled based on assumed or statistically represented loads spectrum, which doesn’t necessarily fit to the way which a given aircraft is operated. This jeopardize the safety, but also is connected with not scheduled inspections, whose costs are much higher than regular ones. In fact, the fraction of unexpected NDI prevail over scheduled inspections. Furthermore, application of new lightweight materials, e.g. composites, introduces new damage evolution pathways, making it difficult to use low cost NDT techniques like visual testing, which account for about 60% of overall NDI. Therefore, there is a strong need from the industry sector to introduce Structural Health Monitoring (SHM) and Operational Load Monitoring (OLM) systems, based on sensors permanently integrated with the aircraft structure. Application of such systems would definitely increase safety, especially when considering hardly accessible ‘hot-spots’, but it could also save up to 50% of necessary inspections time depending on the aircraft type. Furthermore, possessing the knowledge about the current state of an aircraft as well as the way it is used, would allow to predict its further performance and determine the optimal time for its overhauls. Clearly, damage detection and damage assessment capabilities of SHM systems, e.g. expressed in terms of PoD curves, are the most important ones. However, from the cost analysis perspective, also a very important property of SHM systems is their false calls ratio. Damage indication by SHM system will be verified with classical NDT methods, at least at their early stage of development, which would increase the number of unplanned NDI and could rise aircraft maintenance costs, if there would be too much false positive findings. Both, the improvement of damage detection capabilities as well as the reduction of false calls ratio of SHM systems, are as much important for their applicability.Providing reliable and universal Structural Health Monitoring (SHM) system allowing for direct aircraft inspections and maintenance costs reduction is one of the major issues in the aerospace industry. The installation of SHM sensors in composite structures for structural health monitoring requires care selection of a proper techniques to guarantee its reliability and lack of affecting structure durability. There are several SHM technologies which are exist based on the principles of the Non Destructive Evaluation Technology. Among them we may differentiate: techniques based on guided Lamb waves, acoustic emission, strain monitoring based on Fiber Optic Sensing.In the paper an approach for the implementation to military aerospace structures damage growth monitoring and early damage detection of the structural elements is presented based on selected SHM techniques. Advantages and limitations of addressed technologies are presented. Experience of the authors with described technologies to composite structures are highlighted. In particular some issues concerning the mathematical algorithms inferring about damage from the impact damage presence and its growth are discussed. Further perspective of the SHM implementation on military aerial platforms based on lessons learned will be also delivered.

Krzysztof Dragan, Michał Dziendzikowski, Artur Kurnyta, Kamil Kowalczyk

Real-Time Stress Concentration Monitoring of Aircraft Structure During Flights Using Optical Fiber Distributed Sensor with High Spatial Resolution

For efficient and reliable design and operation of aircraft, monitoring of deformation and strain during flights is beneficial. Stress concentration is specifically critical for structural integrity, whereas it is a difficult phenomenon to predict and accurately observe. Optical fiber sensors are highly applicable for this purpose because they are light-weighted and flexible, which allows less invasive installation to aircraft structures. In addition, they are capable of distributed monitoring, which enables efficient measurements by collecting strain distributions along a single fiber. We have developed an optical fiber distributed sensing technique with a high spatial resolution which allowed us to monitor strain distribution profiles and stress concentrations. We use long-length fiber Bragg gratings (FBGs) and optical frequency domain reflectometry (OFDR). The OFDR-FBG technique allows us to monitor real-time strain distributions during flights along the FBGs with a 1.6 mm spatial resolution and a 151 Hz sampling rate. We applied this monitoring technique to a middle-sized flying test bed, and conducted flight tests. We measured strain distributions of a fuselage stringer, an aft-bulkhead and a main wing during flights. We report time histories of strain distribution profiles during various maneuvers especially focusing on stress concentration areas. We also show a record of accumulated strain experiences of the wing, which is expected to be an essential feedback for fatigue analysis.

Daichi Wada, Hirotaka Igawa, Masato Tamayama, Tokio Kasai, Hitoshi Arizono, Hideaki Murayama

Research on the Scatter of Structural Load-Time History in a Fleet

Scatter of structural fatigue life is mainly caused by the variability in structures and load-time histories. To study the scatter of structural load-time history of aircrafts in a fleet, this research is focused on analysis of operational load spectra at four typical structure details, of which representative coupon fatigue tests have been carried out. First, operational load-time histories at critical locations in question of each aircraft are obtained according to the flight-data based load equation and transfer function relating the monitored load to the critical location stresses. Then, the fatigue notch coefficient in local stress-strain method is calibrated and validated by representative coupon test data and then used to calculate fatigue damage under different structural load-time histories of each aircraft. Finally, statistical analysis of the fatigue damages is conducted on the assumption of log-normal distribution and scatter factor are obtained. It has indicated that there are significant differences in the scatter factor of different structural load-time histories. It may be advisable to fully consider the differences of scatter factor among different structures in structure fatigue life assessment.

Tang Li, Yongjun Wang, Hongna Dui, Jiang Dong

Study of Composite Impact Dent Visual Detectability and Damage Relaxation Phenomena

The studies focused on visual detectability of surface damages in aircraft composite skins were performed. The goals were to understand how inspection conditions and service factors affect the detectability during standard field control procedures, to establish reasonable barely visible impact damage threshold and to optimize maintenance program in the frameworks of Irkut MC-21 aircraft certification. The probability of surface damage detection as a function of damage size was experimentally evaluated on empennage-type stringer panels in relation to qualification of experts, paint color, viewing distance and surface contamination. The significance of those factors was estimated by non-parametric methods of mathematical statistics, the bootstrap technique was applied for empirical data reduction. Based on the criteria, accepted in industry, the thresholds of detectability were established for general and detailed visual inspection procedures. The special attention was paid to the investigation of impact dent relaxation phenomena, i.e. to the reduction of damage size in time. The relaxation tests were conducted under normal and under hot-wet conditions and it was determined that combination of moisture and elevated temperature leads to the maximum damage size reduction. The performed studies resulted in recommendations on actual external damage size values which should be used as input data in damage tolerance analysis and as detectability criteria in aircraft maintenance manual.

Stanislav Dubinskii, Vitaliy Senik, Yuri Feygenbaum

Study of Load Spectrum Occurring in the Course of Photogrammetric Missions of the UAV

Among common applications of light-weight civilian UAVs are aerial monitoring and photogrammetry. A good example of such UAVs’ application is MONICA Project, devoted to aerial monitoring and photogrammetry of Antarctic Specially Protected Areas (ASPA) on King George Island (KGI). It involved the PW-ZOOM used as a UAV platform, a plane powered by 4hp gas engine, having 3.2 m of wing span and 24 kg of take-off weight. During three Antarctic spring seasons (2014–2016) the PW-ZOOM performed 33 flights of total distance of 3641 km, and spent 35,6 h in the air over the coastal zone of KGI. Besides delivering a large amount of photogrammetry data, also the flight-logs with several flights parameters were collected, which allowed for studying various flight dynamic aspects.The paper contains an analysis of the flight-logs, with a focus on investigation of load spectrum during photogrammetry missions. The analysis was based on flights having the same scenario and flight-track, but performed in different weather conditions, especially at different wind speeds. The load spectra were developed as half-cycle arrays of load factor signal (i.e. transfer arrays based on the rainflow counting algorithm), and as incremental load spectra. Calculation algorithms were prepared by the authors in the LabVIEW environment (Fig. 1). The obtained results allow for assessing the influence of weather on the fatigue loads during flight. They can be useful as the basis necessary for load spectrum extrapolation and preparation of fatigue tests of the PW-ZOOM structure or similar planes. Fig. 1. The PW-ZOOM – an UAV for photogrammetry missions.

Miroslaw Rodzewicz, Dominik Glowacki

Substitute Models for Structural Components Loads Estimation Based on Flight Parameters and Statistical Inference Methods

The knowledge about loads of the aircraft structure occurring during aircraft operations, is one of the fundamental elements of not only damage tolerance approach to aircraft design. Nowadays Operational Loads Monitoring (OLM) programs are essential elements of airworthiness policy of Ministry of Defense in some NATO countries. In the OLM case, aircraft loads related information could be available from a sensors network e.g. strain gauges or even Fiber Bragg Gratings (FBG) permanently mounted in the aircraft structure and measuring its local strains (direct loads monitoring). One of the key issues for direct loads monitoring, is to identify different distribution of stresses occurring in the structure, in order to apply appropriate models for prediction of damage evolution. When identified, number of load cycles of each type can be determined and their contribution to fatigue can be calculated.Due to complexity and costs of OLM programs with direct loads monitoring, indirect loads monitoring approach, based on flight parameters analysis, has lost nothing of its attractiveness. What is available instead of local strains distribution, is a set of recorded flight parameters, which by the laws of inertia and aerodynamics should determine dominant part of loads, acting on a given element. Compared to direct load monitoring, an additional difficulty emerges. Beside identification of different stress distribution, also functional relation between stress field and flight parameters needs to be established. In the paper, Canonical Correlation Analysis (CCA) method is discussed as an useful method for selection of flight parameters which can properly predict aerodynamic loads acting on a given structure. CCA allows both for identification of different modes of stress distribution as well as identification of flight parameters which are the best suited for their prediction.

Michal Dziendzikowski, Wojciech Zielinski, Piotr Reymer, Marcin Kurdelski, Piotr Synaszko, Witold Klimczyk, Andrzej Leski, Krzysztof Dragan

Technical Justification for an Ultrasonic Inspection Procedure Applied to a Helicopter Component

For many non-destructive evaluation (NDE) applications, traditional probability of detection (POD) assessments are impractical because of the cost, time, and complexity associated with manufacturing and preparing the required specimens representative of in-service conditions. Various alternative methods have been developed to reduce the number of test specimens required for the reliability estimation.Technical justification is a process that includes analytical and experimental evidences, physical reasoning, summary and recommendation. Those are gathered and compiled in a structured format to verify that the targeted inspection technique, equipment and written procedure conform to the requirements and can meet its stated objective. Inspection qualification through technical justification minimizes the reliance on the manufacturing of test pieces and their time consuming inspection trials. In this paper, this promising approach is applied to demonstrate the reliability of an ultrasonic NDE procedure for the inspection of a helicopter upper tail-cone assembly. It is aimed to provide comprehensive evidence for determining whether the minimum detectable discontinuity size by the existing ultrasonic inspection procedure can be reduced from 1.27 mm (0.050 in.) to 0.64 mm (0.025 in.), without compromising the current level of confidence.

Muzibur Khan

The Research of Aircraft Structure Health Monitoring System Based on Big Data Analysis

The aircraft Structure Health Monitoring System (SHMS) is an important system for the old age or future aircrafts. The main components of this system includes many new and advanced sensors like fiber optic sensor etc. in the principal structure elements of aircraft structures, and the data processing module which needs to collect and demodulate the massive data, and the data analysis and storage which can be used to predict the structure’s life and evaluate the health status of the whole aircraft.This paper introduces the concept and functions of the SHMS, and presents the state-of-the-art for SHMS all-around of China, and the development history of SHM and aircraft life management during the past several decades. As the technology of big data analysis and cloud computing becomes more and more mature, a new research direction has been proposed in this paper.This paper provides the operating principle of SHMS and the flow of big data including the data collection and analysis and management. The system frame design method is proposed, and both of software and hardware conception also be described in this paper. The main functions and maintenance of SHMS also be discussed for the system design.Last but not the least, this paper provides the assumption that the aircraft structure life follows the lognormal distribution and the research on comparison between the aircraft structure life with and without structure health monitoring system. The analysis result shows that the aircraft use life and average life grows as the reliability of structure health monitoring grows and the aircraft structure maintenance become more effective.

Zhinan Zhang, Yu Ning, Xinbo Wang, Bintuan Wang


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