The design of advanced aircraft structures to resist acoustic fatigue; work in progress

Details of the current research activities in the Department of Aeronautics and Astronautics, on the response of advanced aircraft structures to acoustic loading, are presented. Both box type and sandwich structures, employing aluminium alloy, carbon fibre reinforced plastic, and GLARE composite materials, are being investigated. In order to develop design guidelines for these complex structures, it is necessary to combine theoretical predictions, using the finite element method, with experimental measurements of the structural response to random acoustic loading. Both types of structure will be tested in the Progressive Wave Tube facility at Southampton. In addition, it is hoped that a more comprehensive damping guide will be produced for the type of structure used in advanced aircraft design.     

A reliability-based framework for fatigue damage prognosis of composite aircraft structures

The extensive use of lightweight composite materials in composite aircraft structures drastically increases the sensitivity to both fatigue- and impact-induced damage of their critical structural components during their service life. Within this scenario, an integrated hardware–software system that is capable of monitoring the composite airframe, assessing its structural integrity, identifying a condition-based maintenance, and predicting the remaining service life of its critical components is therefore needed. As a contribution to this goal, this paper presents the theoretical basis of a novel and comprehensive probabilistic methodology for predicting the remaining service life of adhesively bonded joints within the structural components of composite aircraft, with emphasis on a composite wing structure. Non-destructive evaluation techniques and recursive Bayesian inference are used to (i) assess the current state of damage of the system and (ii) update the joint probability distribution function (PDF) of the damage extents at various locations. A probabilistic model for future aerodynamic loads and a damage evolution model for the adhesive are then used to stochastically propagate damage through the joints and predict the joint PDF of the damage extents at future times. This information is subsequently used to probabilistically assess the reduced (due to damage) global aeroelastic performance of the wing by computing the PDFs of its flutter velocity and the velocities associated with the limit cycle oscillations of interest. Combined local and global failure criteria are finally used to compute lower and upper bounds for the reliability index of the composite wing structure at future times.     

Environmental effects on fracture resistant and biaxial fatigue design of aircraft structures

This paper discusses the two interrelated fields of crack initiation and crack propagation by presenting analytical techniques for calculating fatigue damage in biaxially stressed structures along with determining safe inspection intervals for contained crack growth.The equivalent stress concept is used to derive a set of uniaxial stresses that produce the same amount of fatigue damage as the biaxial stress exposure. The distortion energy concept serves as the basis for combining alternating principal stresses by translating the uniaxial SN curves for zero mean stress into a family of concentric ellipses. The major axis bisects the principal stress axes for isotropic materials with invariant directional fatigue performance. The intercept of these ellipses with the maximum alternating principal stress axis defines the equivalent alternating stress at various cyclic lives.Empirical procedures are given for treating problems with varying principal stress directions and areas with directional dependent fatigue performance.The Goodman diagram relates alternating stresses and mean stresses at any constant cyclic life. If two uniaxial Goodman diagrams are constructed on each reference axis, a three-dimensional body can be visualized which intercepts the zero alternating stress plane in a shape identical to that described by the applicable static load criterion. The equivalent mean stress concept is based on the existence of a similarly shaped closed surface at any value of alternating stress. The intercept of this surface boundary with the maximum mean stress axis is the equivalent mean stress.Crack growth rates and residual strength of structure are important items since it is necessary to consider the possible existence of cracks. Static failsafety consists of contained crack growth for reasonable lengths of time and back-up structure providing alternate load paths. The stress intensity factor $\text{K}$, reflecting the distribution of stress in cracked structure, is the basis for computing crack growth. Baseline crack growth data for several material toughnesses and environmental exposures is required for fracture analysis. The method employed consists of calculating stress intensities for various crack lengths in the structure, these primarily being a function of geometry and applied stress distribution. The crack growth curves are constructed by integrating the baseline data for the appropriate corrosive exposure with additional factors applied to allow for scatter in growth rates and load magnitudes.     

A unified engineering approach to the prediction of multiaxial fatigue fracture of aircraft structures

In this paper the existing biaxial fatigue theories are reviewed. The effect of isotropy, mean stress, phase angle, and notches on biaxial fatigue is discussed. An approach based on equivalent stress is proposed. The exactness and consistency of this approach is verified with experimental results of full scale test articles. The analysis indicates that this simple approach can be used with confidence in predicting the linear cumulative damage in full scale structural components, which are experiencing multiaxial stress loading.     

An elastic-plastic finite element alternating method for analyzing wide-spread fatigue damage in aircraft structures

In this paper, a new analytical technique to study the effect of wide-spread fatigue damage in ductile panels is presented. The main purpose of the study is to develop an efficient methodology to predict the maximum load carrying capacity of panels with cracks. The problem arises especially in the fuselage skin of aging airplanes, in which cracks initiate from a row of rivet holes. This problem is known as Multi Site Damage (MSD) in aging aircraft. It is very important to estimate the load carrying capacity. Usually, the approach based on elastic fracture mechanics may overestimate the load capacity. It is very important for the aircraft structure with MSD to estimate the load carrying capacity of such damaged structures. Approaches based on elastic fracture mechanics often lead to a considerable error. In this paper, the Elastic Finite Element Alternating Method (EFEAM) has been extended to the case of elastic-plastic fracture of panels with MSD cracks. In EFEAM, analytical solutions to crack problems in an infinite plate are employed. In this study, we adopted an analytical solution for a row of cracks in an infinite panel. Furthermore, the plastic deformation is accounted for, by using the initial stress algorithm. The T _inf ^sup* integral is employed for the fracture criterion. The methodology developed in the present study can be called as Elastic-Plastic Finite Element Alternating Method (EPFEAM) for MSD problems. A series of studies on the maximum load capacity of panels with a row of cracks has been conducted.The support of this work by the Federal Aviation Administration through a grant to the Center of Excellence for Computational Modeling of Aircraft Structures, at the Georgia Institute of Technology, is sincerely appreciated.     

Approximate determination of a strain-controlled fatigue life curve for aluminum alloy sheets for aircraft structures

This paper deals with selected methods of approximate determination of a strain-controlled fatigue life curve for aluminum alloy sheets used in aircraft structures. Authors based their analysis of those methods on the results of own research of 2024-T3 alloy and its Russian equivalent D16CzATW. The approximate strain-fatigue life curves were compared with the experimental curves. The influence of inconsistencies between those curves on the calculation results was analyzed on computational examples by means of the Palmgren–Miner’s rule.     

Direct measurement of fatigue damage in aircraft

A fatigue life gauge, mounted on a mechanical amplifier and bonded to a structure, will integrate the load/frequency pattern to which it is exposed. Subject to static calibration establishing the strain/load relationship registered by the device, gauge manufacturers data will closely predict the output for programmed loading within the range of 0–2 to 40 ohms. The purpose of the device is to compare the rate of loading, accumulated by a test structure, with the rate accruing to similar structures undergoing variable service loading. Inasmuch as this loading affects the fatigue life of a component, the device will monitor fatigue damage. Some agreement is shown between laboratory test of an aircraft fin and measurements taken during routine flying. The device is sensitive to change in aircraft utilisation. In the case of fin structure, no correlation exists between the measured damage factor and the damage assessed by the standard aircraft fatigue meter.     

A new conception for enhancement of fatigue life of large number of fastener holes in aircraft structures

A new conception for increasing fatigue life of large number of fastener holes in aircraft structures is developed. It is accomplished by a new method, called friction stir hole expansion (FSHE). This method not only reduces labour and time consumption, but it also decreases the overall cost for processing a large number of holes in structures made of aerospace grade 2024‐T3 aluminium alloy. FSHE combines the advantages of friction stir processing with these of mandrel cold working methods in two ways: a micro effect, expressed in hole surface modification, and a macro effect, expressed by the introduction of beneficial compressive residual macro stresses. The effectiveness of the method has been assessed by fatigue tests. Finite element simulations have been carried out. It has been proven that the greater fatigue life of fastener holes, processed by FSHE, is a consequence of the obtained micro effect.     

Some experiences in aircraft structures

When considering the Chairman's distinguished invitation to present a talk on a technical subject of my choice, I have come to the conclusions that a historical survey of the Dutch activities in structures might be appropriate, in particular of the older history.     

Investigation of the fatigue and long-term strength of flat duralumin elements of aircraft structures in certain aggressive media

The effect of various aggressive media encountered in agricultural aircraft applications on the fatigue and long-term strength of flat riveted duralumin specimens (simulating the casing of various aircraft parts) was studied. It was shown that crevice corrosion in the zone of contact between various parts leads to a substantial reduction in their fatigue and long-term strength.     

Damage tolerance of bonded aircraft structures

This paper presents a damage tolerance philosophy for bonded structures and repairs. It is proposed to assess the damage growth in bonded structures loaded mainly in shear with a generic strain elastic energy release rate concept. This concept has been validated on metal-to-metal and metal-to-composite bonding in metallic and hybrid structures.     

Generation of correlated stress time histories from continuous turbulence Power Spectral Density for fatigue analysis of aircraft structures

The main purpose of the present work is to develop an efficient computational method for generating correlated stress time histories for aircraft structures under gust loads. Random gusts in any direction, which lead to random multiaxial loads on the aircraft structure are considered. A direct temporal simulation of stress using a finite element model is not possible due to computational burden. In order to overcome this, a new and efficient method based on Power Spectral Density functions (PSD) of stress is proposed. Since the PSD of the various stress components have not previously been correlated, a result enabling the direct generation of the correlation between them has been established, which is crucial for fatigue and damage tolerance analysis in several dimensions. Validation is performed by comparison with a direct (and costly) temporal simulation as well as Rice’s formulas. An example on a long range aircraft illustrates the relevance of the proposed approach. Although this method is presented with an application on a two dimensional stress tensor, it should be noted that it could be straightforwardly extended to any linear system with multiple input and output.     

Review of fatigue monitoring of agile military aircraft

Fatigue monitoring of airframes has developed over the decades to the stage where it is now incumbent for all fighter type aircraft to be fitted with an airborne fatigue monitoring system. These systems typically collect operational data for the calculation of the safe-life or the inspection interval of the airframe.
This paper presents a state-of-the-art review of fatigue monitoring systems of agile military aircraft. It reviews and comprehensively examines the techniques used in individual aircraft fatigue monitoring programs, and examines current systems and practises. Based on experience from Australian fatigue monitoring programs, it highlights some of the potential pitfalls in the systems and techniques. It also investigates the issues of strain gauge utilization and calibration, collection of flight parameter data, data integrity, comparisons with fatigue test results and fatigue damage models. Some of the problems with current systems are highlighted and requirements for future fatigue monitoring systems are suggested.
This review has determined that there is little uniformity in the fatigue management practices of operators and that many aspects of the fatigue management process have been overlooked by some structural integrity managers. Also, very few of the papers reviewed specified the philosophy or aims of their monitoring systems.     

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.     

Service life and survivability of aircraft structures

We present a methodology aimed at guaranteeing high fatigue and crack-growth resistance of aircraft structures in Russia and formulate criteria of survivability of these structures and requirements to the materials intended for the use in designed structures. We also analyze the methods used for the solution of the problem of safe operation of old aircrafts which take into account the degradation of fatigue and crack-growth resistance of structural materials and suggest a solution of the problem of multiple-origin cracks for aircraft structures in Russia. Central Institute of Aerohydrodynamics (TsAGI) Zhukovskii, 140160, Moscow Region, Russia. Translated from Fizyko-Khimichna Mekhanika Materialiv, Vol. 32, No. 2, pp. 43–56, March–April, 1996.     

Multiple-site and widespread fatigue damage in aging aircraft

Transport is a vital cog in the world economy. However, the cost of new aircraft has meant that there are now aging fleets whose continued airworthiness requires special analysis and improved crack detection techniques. This paper highlights the need for research into the field of structures containing both multi-site and widespread fatigue damage, topics which have recently received worldwide attention. To adequately explain the work that has been completed in these fields, a brief introduction to some of the more common terminology will also be given.     

A survey of serious aircraft accidents involving fatigue fracture

A world-wide survey of serious aircraft accidents involving fatigue fracture has been carried out. The study includes not only fatal accidents, but also those in which the damage to the airframe was substantial or greater. The accidents cover civil and, to a limited extent, military aircraft. A total of 1885 accidents since 1927 were identified as having fatigue fracture as a related cause, and these accidents resulted in 2240 deaths. Engine/transmission failure and landing-gear failure were the most common cause of recent fixed-wing accidents, while the most prevalent rotary-wing problems were failure of the engine/transmission and of the tail-rotor. Currently there is a yearly average of about 100 serious fatigue accidents (69 fixed-wing and 31 rotary-wing).     

Typical fatigue-initiating discontinuities in metallic aircraft structures

A fatigue lifing framework using a lead crack concept, based on years of detailed inspection and analysis of fatigue cracks in many specimens and airframe components, has been developed by the DSTO for metallic primary airframe components. This framework is an important additional tool for determining aircraft component fatigue lives in the Royal Australian Air Force fleet. Like the original Damage Tolerance concept, developed by the United States Air Force, this framework assumes that fatigue cracking begins as soon as an aircraft enters service. However, there are major and fundamental differences. Instead of assuming initial crack sizes and deriving early crack growth behaviour from back-extrapolation of growth data for long cracks, the framework uses data for real cracks growing from small discontinuities inherent to the material and the production of the component. To this end, this paper examines the types of discontinuities that initiate fatigue cracks in typical metallic airframe structures. These discontinuities and the fatigue cracks that have grown from them are taken from coupon, component and full-scale tests, and also from service aircraft, including commercial transport aircraft and high performance military aircraft.