Crack Paths 2009

Analysis of spectrum fatigue crack growth in AA7050

specimens with three stress concentrations

W.Zhuangand L. Molent

Air Vehicles Division, Defence Science and Technology Organisation, Melbourne, VIC

3207, Australia. wyman.zhuang@defence.gov.au

ABSTRACT.Fatigue crack growth behavior in aluminium alloy 7050-T7451 under a

fighter aircraft wing root bending moment spectrum was investigated. The crack growth

data were measured by quantitative fractography for three groups of specimens with

different stress concentration geometrical features. Based on the analysis of the

measured spectrum crack growth data using linear elastic fracture mechanics, it was

found that the concept of geometry factors in the stress intensity factor could not

collapse the crack growth rate data derived from each stress concentration feature,

particularly near the small crack growth region. In order to investigate the possible

reasons for this, finite element analysis was used to determine notch plastic zone sizes

for each stress concentration geometry. As a consequence, an alternative crack growth

driving force by considering both notch elastic-plastic stress field and gross net-section

stress field was proposed and used to interpret the fatigue crack growth data under

spectrum loading.

I N T R O D U C T I O N

Aircraft structures exposed to variable amplitude flight spectrum loading are prone to

2]. Fatigue critical structures in aircraft also contain a range of

fatigue failure [1,

geometric discontinuities such as fastener holes and cutouts that can cause stress

concentrations or create hot spots for fatigue crack initiation and propagation [3].

Therefore, the objective of this study is to investigate fatigue crack growth (FCG)

behavior in aluminium alloy specimens with different stress concentration features,

subjected to variable amplitude loading.

An extensive research effort has been devoted to investigating the effect of geometric

discontinuities, namely notches on F C Gbehavior due to their importance to aircraft

structural damage tolerance analysis. It is often required to predict the fatigue life of

notched members based on F C Gdata obtained experimentally from the specimens or

structures with different stress concentration factors ( ) or even smooth specimens [ t K 4,

5, 6, 7, 8].

It is noteworthy that all these studies were carried based on constant

amplitude (CA) fatigue tests at different stress ratios . For instance, Figure 1 shows the F C Gbehavior for diff rent notch root radii ( R ) in a low carbon steel plate with a U

shaped notch under C A loading at

[8]. It demonstrated that the crack growth rate

R = 0

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