Damage tolerance demonstration testing for the Australian F/A-18

Over recent years Australia has been involved in a number of full-scale fatigue testing programs in support of the through-life structural integrity of the Royal Australian Air Force’s (RAAF) F/A-18 fleet. It was recognised early in the acquisition cycle that the certification testing conducted by the manufacturer failed to considered damage tolerance requirements and would be unlikely to cover the typically more severe and diverse RAAF operations. Given similar aircraft structural integrity management philosophies, major benefits were to be realised through collaboration with the Canadian Forces (CF). In particular, as fatigue testing under representative CF/RAAF loading was the basis for both countries’ structural integrity management, the International Follow-On Structural Test Project (IFOSTP) was successfully concluded.This paper emphasises the Australian components of IFOSTP, including the damage tolerance testing and demonstration for the aft fuselage that incorporated the simultaneous application of both manoeuvre and dynamic buffet loads. Many of the innovations and consequences of this work program are highlighted, and may be applicable to future fighter aircraft structural integrity programs.     

A review of some damage tolerance design approaches for aircraft structures

Damage tolerance design is becoming a necessity in the design of modern aircraft although its importance was recognized as long as four centuries ago by Leonardo da Vinci. Two decades ago structural design engineers and research workers felt the need of incorporating damage tolerance in the design of aircraft structure. Due to a lack of comprehensive damage tolerance methodology large scale component test results were used to develop empirical damage tolerance methods. Recently, linear elastic fracture mechanics has been used in predicting residual strength and crack growth rates in damaged structure. As a result of these efforts significant developments in cracked structure analytical methodology have been achieved. The recent Air Force requirement to apply linear elastic fracture mechanics approach in damage tolerance design of aircraft structures, warrants and critical review of various approaches. In this paper an attempt has been made to critically review some damage tolerance design approaches and their application to aircraft structures.

The paper consists of three main sections: The first section reviews the residual strength analysis methodology, assumptions and limitations of each method are discussed through a simple example. The second part surveys the various crack propagation laws, including linear and non-linear ranges and spectrum loading effects. In the third and last section, fracture mechanics methodology is applied to several types of built-up structural components under spectrum loading conditions. The comparison of test results and analysis of complex structures indicate that simple methods of fracture mechanics can be applied to find the damage tolerant strength and rate of crack growth.

The review presented in this paper indicates that the majority of work done in development of fracture mechanics analytical methodology has been based on data obtained from small scale laboratory specimens tested under closely controlled conditions of damage and environment. The validity of the methodology for complex structure under complex loading conditions has not been established. Before the results of a fracture mechanics analytical methodology can be accepted with a high degree of confidence many realism factors must be properly accounted for in the analysis.     

Materials Mechanics – The Basis of Advanced Technology for Ageing Aircraft

Materials mechanics plays a significant role in assessing the damage status and handling of an ageing aircraft. This paper deals with determination and use of operational loads monitoring regarding fatigue life evaluation at specifically damage critical locations as well as the resulting consequences in fleet management. Regarding damage‐tolerant design specifically aspects such as simulation of multi‐site damage cracking and optimised design of repairs are discussed. Finally some recent techniques and results with regard to structure integrated sensing in the context of smart structures is presented.     

基于载荷损伤分散的严重谱选取方法初探

在飞机结构设计定型阶段,要综合考虑结构的差异和载荷的分散性评定机群的安全寿命。美国空海军联合规范指南JSSG-2006、标准MIL-STD-1530C和我国国军标GJB67.6A-2008提出采用90%谱(严重谱)进行飞机结构耐久性分析和试验,但并未给出严重谱严重程度的选取准则。该文以机群载荷损伤分散描述机群载荷分散,初步探讨了载荷谱严重程度的选取方法。假定指定谱下的结构寿命和机群载荷损伤均服从对数正态分布,在综合载荷损伤和结构分散性的寿命分布基础上,分析了严重谱下的寿命可靠度。按机群安全寿命满足99.9%的可靠度要求,以载荷损伤覆盖概率描述载荷谱严重程度,给出了载荷谱严重程度的表示形式,并针对典型的飞机结构分散和载荷损伤分散性参数,给出了典型的载荷损伤覆盖概率取值。分析表明严重谱的严重程度与载荷损伤分散性和结构分散性均有关,基于损伤的90%严重谱可以保证机群内超过99%飞机的使用安全。     

On damage tolerance design of fuselage structure (longitudinal cracks)

Aircraft structure is the most obvious example where functional requirements demand light weight and, therefore, high operating stresses. An efficient structural component must have three primary attributes; namely, the ability to perform its intended function, adequate service life and the capability of being produced at reasonable cost. To ensure the safety of aircraft structures, the Air Force requires damage tolerance analysis. This paper focuses its attention on designing a fail-safe fuselage structure. Two types of damage most frequently associated with the structural integrity of the fuselage are longitudinal cracks under high hoop stresses induced by cabin pressurization and circumferential cracks under stresses from vertical bending of the fuselage. The analysis of these types of cracks is complex, first due to the complex structural configuration (i.e. frames, skin longeron and crack stopper straps) and secondly due to the influence of the curvature of the shell. Various analytical and empirical approaches have been used to study the damage tolerance capability of the fuselage structure. Due to the lack of a comprehensive model to calculate the stress intensity factors for the complex structure, experiments usually have been performed to measure the crack growth rates and to demonstrate the residual strength of fuselage-type structural components containing circumferential and longitudinal cracks.

In this paper various analytical and empirical approaches used in evaluating the damage tolerance capability of the fuselage structure are critically evaluated and compared. A model which accounts for the influence of frames, straps and curvature is developed. This model is then used in an example problem having typical military cargo aircraft fuselage structural elements. The Air Force damage tolerance requirements are discussed in detail.     

基于阈值超越概率的结构耐久性与损伤容限综合分析方法

飞机结构耐久性分析和损伤容限分析都是以裂纹扩展分析为基础的,该文根据二者之间的内在联系,将耐久性分析与损伤容限分析进行综合考虑。该方法以损伤容限设计手册中规定的非概率初始裂纹尺寸作为阈值裂纹尺寸,根据裂纹尺寸超过阈值的概率以及耐久性分析结果确定概率损伤容限分析时的初始裂纹尺寸分布。裂纹尺寸超过阈值的概率可以通过检测数据给出,评估结果较为可靠。在概率损伤容限分析时,为了避免多重积分的计算,给出了结构安全余量的显式表达式以及一次二阶矩方法的计算公式,提高了可靠度计算效率。该文结合工程实例并与现有方法进行了对比分析,结果表明该方法是合理的、有效的。     

Verification of an individual aircraft fatigue monitoring system

Fatigue monitoring of airframes has developed over the decades to the stage where it is now incumbent for the certification of fighter type aircraft to incorporate a fatigue monitoring system. These systems typically collect operational data for the calculation of the airframe’s safe-life or crack inspection intervals. Many of these systems are complex, incorporating such features as data integrity checking, strain gauge calibration algorithms and damage calculation algorithms to name a few. Whilst it may be possible to validate the robustness and accuracy of specific system components (e.g. the damage algorithm can be tested against fatigue coupon results), the verification of the performance of the in-service system as a whole presents a much bigger challenge.In this paper, the verification of the Royal Australian Air Force’s F/A-18A/B Hornet individual aircraft fatigue monitoring system is outlined. The availability of a significant number of ex-service centre fuselage sections with known usage has facilitated this effort. Using an enhanced teardown procedure, in-service fatigue crack growth has been identified at a significant number of locations. All the in-service cracking corresponded to the same locations found cracked in the fatigue certification full-scale test article that was used to calibrate the usage monitoring system, so that by comparing the measured in-service growth with the test-demonstrated growth the functionality of the monitoring system could be assessed. This assessment should reveal the effectiveness of the system in providing robust fatigue life expended indices to help ensure that structural integrity boundaries are not exceeded. For this comparison, the crack growth was measured using quantitative fractography.It is believed that this work is the first example of using the crack growth in retired structure of known usage to verify a fatigue tracking system that incorporated significant aircraft structural integrity elements including tracking philosophy, structural fatigue lifing methodology, full-scale fatigue test results, design standard interpretation and retirement considerations.     

Fatigue issues in aircraft maintenance and repairs

Many design considerations are involved in ensuring structural integrity of Boeing jet transports, which have common design features validated by extensive analyses, tests, and service performance. Designing for continued structural integrity in the presence of damage such as fatigue or corrosion is an evolutionary process. Performance demands, increasing structural complexity, and aging fleet reassessments have required development of standards suitable for application by large teams of engineers. This presentation is focused on such methods with special emphasis on practical fatigue reliability considerations. Durability evaluations are based on quantitative structural fatigue ratings which incorporate reliability considerations for test data reduction and fleet performance predictions. Fatigue damage detection assessments are based on detection reliability estimates coupled to damage growth and residual strength evaluations. Data are presented to airline operators on detection check forms which permit efficient maintenance planning to achieve required fatigue damage detection reliability levels.     

Fatigue crack growth and damage tolerance

The problems arising as a result of aging aircraft, rail and civil infrastructure have focused attention on tools for predicting the growth of cracks from small naturally occurring material discontinuities. To this end, the present paper discusses on the difference between the analysis tools needed for ab initio design and sustainment, modelling of cracks that grow from small naturally occurring material discontinuities and ways to determine the short crack da/dN versus ΔK data from long crack American Society for Testing and Materials (ASTM) tests. It also discusses how existing equations can be used to predict short crack growth and how to account for the variations seen in crack growth histories. Attention is also focused on the recent Federal Aviation Administration limit of validity ruling and the effect of the environment on widespread fatigue damage in civil transport aircraft.     

Probability modelling and statistical analysis of damage in the lower wing skins of two retired B-707 aircraft

A plausible mechanistically based probability model for localized pitting corrosion and subsequent fatigue crack nucleation and growth is used to analyse tear-down inspection data from two retired B-707 aircraft that had been in commercial service for about 24 and 30 years. Sections of the left-hand lower wing skins from these aircraft had been previously disassembled and inspected optically at 20× magnification. The inspections were augmented by metallographic examinations for the lower time aircraft. The evolution of damage in the fastener holes is estimated by using reasonable values for the localized corrosion and fatigue crack growth rates, statistically estimated from laboratory data. The primary loading, assumed to be the mean design load, is considered to be from ground–air–ground wing bending cycles, augmented by 'average' gust loading, only. The encouraging agreement between the estimated probability of occurrence and the observed distribution of multiple hole–wall cracks attests to the efficacy of the approach and its relevancy to airworthiness assessment and fleet life management.     

Methodology for fatigue life assessment of the structural integrity of fighter aircraft

This work describes the present methodology used on Portuguese Air Force (PoAF) aircrafts for fatigue life assessment to maintain its structural integrity. This methodology uses experimental flight data collected by airborne instrumentation systems. The recorded data are analysed for the prediction of the crack growth of defects at critical components of the aircraft and this information is then used for planning non‐destructive inspections, maintenance, logistical support and operational usage of the various fleets. This methodology is exemplified by one application on DASSAULT/DORNIER ALPHA‐JET aircraft.     

On damage tolerance design of fuselage structure—circumferential cracks

The aircraft structure is the most obvious example where functional requirements demand light weight and, therefore, high operating stresses. An efficient structural component must have three primary attributes; namely, the ability to perform its intended function, adequate service life, and the capability of being produced at reasonable cost. To ensure the safety of aircraft structures, the Air Force requires damage tolerance analysis. This paper focuses its attention on designing a fail-safe fuselage structure considering circumferential cracks under stresses from vertical bending of the fuselage. The analysis of these types of cracks is complex, first due to the complex structural configuration (i.e. frames, skin longeron and crack stopper straps) and secondly due to the influence of the curvature of the shell. Various analytical and empirical approaches have been used to study the damage tolerance capability of the fuselage structure. Due to lack of a comprehensive model to calculate the stress intensity factors for the complex structure, experiments usually have been performed to measure the crack growth rates and to demonstrate the residual strength of fuselage-type of structural components containing circumferential and longitudinal cracks. In this paper various analytical and empirical approaches used in evaluating the damage tolerance capability of the fuselage structure are critically evaluated and compared. A model which accounts for the influence of frames, straps and curvature is developed. This model is then used in an example problem having typical military cargo aircraft fuselage structural elements. The Air Force damage tolerance requirements are discussed in detail.     

Block-by-block approaches for spectrum fatigue crack growth prediction

Two block-by-block approaches for improving spectrum fatigue crack growth prediction were proposed and developed in this paper from the observations and analyses of fatigue crack growth behaviours in either representative specimens or real aircraft structures under flight spectrum loading by using the quantitative fractography method. The first approach is the flight-by-flight approach that can be used to predict crack growth history curves for a tested spectrum crack growth data at different stress level for a critical location. The second approach called the effective block approach can be used to predict crack growth histories for un-tested spectra based on some previously tested spectrum crack growth data. In order to demonstrate the robustness of the block-by-block approaches for aircraft damage tolerance analysis, verification and consistency studies were conducted and presented using fatigue test results for different aircraft structures under several flight spectra. It was found that the block-by-block approaches are able to provide significant advantages over conventional fatigue lifing approaches for aircraft damage tolerance analysis.     

Optical fibre sensors for health monitoring of bonded repair systems

In the aircraft industry the use of externally bonded composite repairs has become an accepted way of repairing fatigue, or corrosion, damaged metallic structural components. However, current NDI and damage assessment techniques for composite repairs are passive and generally performed on ground. The challenge is to develop new techniques utilising recent analytical and experimental tools. This report examines the use of optical fibre sensors. Optical fibres offer a means of monitoring the load transfer process in these repairs, and can therefore be used to provide an indication of the integrity of the repair. This paper describes the use of an array of fibre Bragg grating strain sensors (FBGs) for the in situ monitoring of bonded repairs to aircraft structures and, in particular, the monitoring of crack propagation beneath a repair. In this work the FBGs have been multiplexed using a combination of wavelength and spatial techniques employing a tunable Fabry–Pérot (FP) filter to track individual gratings. The multiplexed FBGs were then surface-mounted on a boron–epoxy unidirectional composite patch bonded to an aluminium component. The sensors were located so as to monitor the changing stress field associated with the propagation of a crack beneath the patch. The ability of relating experimental results to sensor readings is then confirmed using both a thermo-elastic scan of the patch and 3D finite element analysis.     

Critical stress intensity factors in steel cracked wires

In the conceptual framework of fracture mechanics analyses, the study of cracked wires axially loaded has the highest interest since numerous structural elements (e.g. wires, cables, cordons or tendons) work under such a type of loading during their service lives. So, a method that allows the determination of stress states at the crack front should be welcome as a useful way for ensuring the structural integrity of those components for different environmental conditions (air, stress corrosion cracking, hydrogen embrittlement,…). To fill this gap, an engineering estimation of the critical stress intensity factor (SIF) is proposed in this paper for eutectoid steel cracked wires under axial loading. The critical SIF is calculated by considering, apart from the fatigue precrack, the subcritical crack propagation before final fracture. Such a subcritical crack propagation is the process zone (by micro-void coalescence MVC) in the case of fracture in air, the subcritical cracking by localized anodic dissolution (LAD) in stress corrosion cracking (SCC) and the tearing topography surface (TTS) in hydrogen assisted cracking (HAC). In addition, different SIF solutions are used in the analysis so as to have a more complete picture of the different phenomena leading to failure and to provide the designer with sound scientific tools. This method allows the engineer to design in the framework of structural integrity and damage tolerance.     

Numerical modelling of crack growth profiles in integral skin-stringer panels

Integral (monolithic) structures can play a significant role in high efficiency structural design. According to the current technological manufacture methods, integral structures have an impact on fabrication cost and on weight reduction. However, until now, some critical aspects have limited the use of these structures. Conventional structures with mechanical or chemical (by adhesion) joints are advantageous because of their damage tolerant and fail safe behaviour. The presence of two separate parts, skins and stringers, guarantee the structural integrity of the component when propagating defects and cracks are present and are thus the key factor in aircraft structures. Focusing our attention on aircraft related structures, the aim of this paper is to show the application of numerical methodologies to evaluate the behaviour of integrally machined skin-stringer panels in the presence of propagating cracks. The described activity resulted from the “Analytical Round Robin on Crack Growth and Residual Strength Prediction in Integral Structures” proposed by ASTM Task Group E08.04.05 while different FE approaches for a single type integral panel with a propagating crack have been introduced in this paper. The crack growth evaluation based on the numerical models agrees well with the experimental results.     

Fracture mechanics modelling of constant and variable amplitude fatigue behaviour of field corroded 7075‐T6511 aluminium

For ageing airframe structures, a critical challenge for next generation linear elastic fracture mechanics (LEFM) modelling is to predict the effect of corrosion damage on the remaining fatigue life and structural integrity of components. This effort aims to extend a previously developed LEFM modelling approach to field corroded specimens and variable amplitude loading. Iterations of LEFM modelling were performed with different initial flaw sizes and crack growth rate laws and compared to detailed experimental measurements of crack formation and small crack growth. Conservative LEFM‐based lifetime predictions of corroded components were achieved using a corrosion modified‐equivalent initial flaw size along with crack growth rates from a constant K_max‐decreasing ΔK protocol. The source of the error in each of the LEFM iterations is critiqued to identify the bounds for engineering application.     

Fatigue damage in aircraft structures,not wanted,but tolerated?

It has been a long way to arrive at the present state of the art about fatigue of structures and materials. Aircraft accidents and incidents were milestones. New concepts were proposed related to structural design, material selection, production techniques, inspection procedures and load spectra. Extensive research efforts have been spent. Our understanding of fatigue damage problems increased significantly. Simultaneously our tools to tackle problems have been developed to a high potential efficiency. And still, there are problems. The present paper is a personal impression of evaluating experience, design aspects, predictions and experiments associated with damage tolerance of aircraft structures.     

Probabilistic and testing analysis for the variability of load spectrum damage in a fleet

Scatter of fatigue life of a fleet is mainly caused by the variability in structures and load spectra. To ensure the safety in service, the probabilistic characterization of load spectrum variability should be researched in durability analysis and testing work. This paper investigates the variability of load damage rate of a fleet. Based on the flight historical parameters measured by individual aircraft tracking (IAT) from hundreds of aircrafts for a certain type of fighter in China, SWT formula and linear damage rule are used to evaluate the load damage, and then, one average and four other individual load spectra are selected corresponding to different damage severities. Fatigue tests are conducted with the Aluminum alloy 7B04-T74 specimens under five spectra and the Titanium alloy TA15M specimens under three of them. The engineering crack initiation lives are measured and the mean lives are estimated assuming the fatigue life following a log-normal distribution. An obvious difference of at least 2.4 times in the load damage rates is found in the fleet. The fatigue lives of a fleet of aircrafts are calculated by Neuber’s approach, and the probabilities refer to damage severities of those 5 load spectra in a fleet are evaluated. The statistical analysis of the fatigue lives and the probabilities shows that a lognormal distribution can be used to describe the variability of load damage rate of a fleet. The variation of the load damage rate is in the same order of magnitude with that in structural properties.