Simon Barter

Marker Loads for Quantitative Fractography of Fatigue Cracks in Aerospace Alloys

The selection of fracture surface marking methods based on exploiting or altering the required fatigue loads is of much interest for many fatigue test programmes. This is particularly true when crack growth measurements during testing are not possible or insufficiently accurate. In such cases, post-test Quantitative Fractography (QF) of the fatigue crack growth may then be needed, and this can be made possible and/or greatly facilitated by fracture surface markers. Here, several examples of fatigue loadings that create fracture surface markings both naturally, as sometimes happens, and intentionally are presented and discussed. While these examples are from fatigue life tests of aircraft alloy specimens and components, particularly high strength aluminium alloys, under normal environmental conditions (air at ambient temperatures), it is probable that some of the fatigue load histories may provide fracture surface markings for other materials and in other environments. The advantages and disadvantages of the various intentional marking methods are detailed with a view to obtaining guidelines and procedures for optimising quantitative fractography of fatigue crack growth. These guidelines are presented in this paper.

Crack growth at low ?K’s and the Frost‐Dugdale law

Abstract This paper examines the fatigue crack growth histories, at low ?K’s, of a range of test specimens and service loaded components and concludes that, as a first approximation, there is a linear relationship between the log of the crack length or depth and the service history (number of load cycles). We also show that, for the cases studied, that the log linear method can give a better prediction of experimental data than a conventional crack growth model.

Aspects of fatigue affecting the design and maintenance of modern military aircraft

Abstract Over recent years, Australian aircraft operators have gained substantial experience with the early detection and management of cracking in critically stressed aircraft components. In the military sphere, the RAAF have been able to operate high-performance aircraft with cracks in specific locations in the airframe, by recognizing that an appropriate safety-by-inspection approach to cracking in critical components could permit continued operation, life extension, and maintenance efficiences. This approach has worked well in older aircraft, where cracks tend to initiate in well-defined high-stress areas, such as at fastener holes or similar design features, and where crack growth is sufficiently localized to allow crack length monitoring using established non-destructive inspection (NDI) techniques. Extension of this approach to newer aircraft, however, presents certain problems; in particular, the use of a variety of life-enhancing treatments (such as shot peening or cold expansion of fastener holes) in combination with high levels of average component stress has led to greater uncertainty about the number of potential cracking sites, their exact location within the structure, and the crack growth rates at these sites once cracks develop. This paper reviews the changing trends in aircraft design with respect to fatigue and fracture resistance, and discusses the influence of factors such as design approach, surface treatment technique, material behaviour and NDI capability on life extension and the management of cracking in a fleet of military aircraft.

Some Practical Implications of Exponential Crack Growth

A review of experimental data show that for many lead fatigue cracks in service components loaded with service spectra, exponential growth (i.e. log crack depth versus cycles or hours) applies for the majority of the life. This behaviour is shown to extend from the micro to macro range of crack sizes in a variety of metals. As a consequence of this, it will also be shown that the crack growth rate is directly proportional to the crack depth. By combining these observations with traditional fracture mechanics approaches to crack growth modelling, a model that is a function of the stress intensity factor (K) with a fixed crack depth influence (non-similitude for the K parameter alone) is proposed. It will then be shown that this model allows for Region I to be smoothly integrated with Region II of the constant amplitude da/dN data. Further, it will be shown that for variable amplitude crack growth data, crack growth ranging from microns to many millimetres can be modelled using this single model. This modelling approach is of particular importance in structural integrity analysis where fatigue cracking cannot always be avoided and the majority of the fatigue life of highly stressed, nominally gross defect free structure is spent growing physically small cracks from initiating discontinuities (i.e. loads in Region I for constant amplitude loading growth rates) up to the point of loss in acceptable strength.

The Effect of Crack Growth Retardation when Comparing Constant Amplitude to Variable Amplitude Loading in an Aluminium Alloy

It has often been observed that the growth of short fatigue cracks under variable amplitude (VA) cyclic loading is not well predicted when utilising standard constant amplitude (CA) crack growth rate/stress intensity data (da/dN v DK). This paper outlines a coupon fatigue test program and analyses, investigating a possible cause of crack growth retardation from CA-only testing. Various test loading spectra were developed with sub-blocks of VA and CA cycles, then using quantitative fractography (QF) the sub-block crack growth increments were measured. Comparison of these results found that, after establishing a consistent uniform crack front using a VA load sequence, the average crack growth rate then progressively slowed down with the number of subsequent CA load cycles applied. Further fractographic investigation of the fracture surface at the end of each CA and VA sub-block crack growth, identified significant crack front morphology differences. Thus it is postulated that a variation or deviation from an efficient crack path is a driver of local retardation in short crack growth during CA loading. This may be a source of error in analytical predictions of crack growth under VA spectra loading that may need to be considered in addition to other potential effects such as less closure whilst cracks are small. For aircraft designers, using solely CA data for fatigue life predictions this may result in non-conservative estimates of total crack fatigue life, producing unexpected failures or an increased maintenance burden.

Metallurgy and Microstructure

Unalloyed titanium has two allotropic forms. The low temperature form, α, exists as an hexagonal-close-packed (hcp) crystal structure up to 882°C, above which it transforms to β, which has a body-centred-cubic (bcc) crystal structure. The alloying behaviour of elements with titanium is defined by their effects on α and β. Element additions that increase or maintain the temperature range of stability of the α phase are called α-stabilizers. The most important of these are aluminium, tin and zirconium. Element additions that stabilize the β phase are called β-stabilizers. These include molybdenum, vanadium and iron. There are also important impurity elements, namely oxygen, hydrogen, nitrogen and carbon. Oxygen and hydrogen are the two most important impurities: oxygen is an α-stabilizer and hydrogen is a β-stabilizer. These four elements are also referred to as interstitial elements. This is because their atomic sizes are much less than those of the metallic alloying elements and they fit in the spaces (interstices) between the crystallographic positions of the metal atoms in the α and β phases.

An Investigation of the Extent of Crack Closure for Crack Growth in an Aluminium Alloy

The corrections incorporated in fatigue crack growth prediction programs for crack closure are usually tested by their ability to predict retardation following an overload and for the accuracy of their prediction lives for long cracks greater than about 1mm. They should, however, be examined on their ability to predict the life of cracks that grow from small sizes, such as small inherent material discontinuities, to failure, which is more typical of service situations and the growth produced by small cycles as well as the larger cycles. To examine the extent of crack closure in aluminium alloy 7050-T7451 and the prediction of that growth, quantitative fractography measurements of short periods of fatigue crack growth produced with a specially engineered spectrum were conducted and are reported here. The spectrum contained bands of constant amplitude loads with diminishing mean stress designed to examine the extent of closure. The quantitative fractography results are compared to predictions by the common analytical programs FASTRAN and AFGROW and further with a basic effective stress intensity calculation method at a crack depth of about 1mm. The results showed that the analytical programs were able to predict the presence of closure; however, the extent of the closure was not accurately predicted.

The quantification of fatigue crack initiators in aluminium alloy 7050‐T7451 using quantitative fractography

Purpose – This paper aims to present some aspects associated with the life prediction of structures with fatigue cracks growing from small natural discontinuities in aluminium alloy (AA)7050‐T7451 for a surface condition that is present in F/A‐18 A/B aircraft critical structure.Design/methodology/approach – Fatigue results are presented for thick section AA7050 plate coupons loaded with a representative fighter aircraft wing root bending moment loading spectrum. Detailed quantitative fractography (QF) was used to gain a deeper understanding of issues relevant to an improved fatigue life predictive capacity for this material by using the QF results to investigate the “effectiveness” of the fatigue initiating discontinuities.Findings – Estimates of the “effectiveness” of the fatigue initiating discontinuities as quasi pre‐existing fatigue cracks (“equivalent pre‐crack size” (EPS) here) were made with the aid of a simple crack growth model. This model, based on experience, was found to be valid for the appli...

Life Extension of F/A-18 LAU-7 Missile Launcher Housings Using Rework Shape Optimisation

LAU-7 missile launcher housings, which are fitted to most Royal Australian Air Force (RAAF) F/A-18A/B aircraft, can experience cracking in the guide rail. This paper covers the design, manufacture and validation of a life extension repair for cracked launcher housings. The repair development uses DSTOs rework shape optimisation technology and fatigue testing capabilities. The rework design reduces peak stresses by 33 %, resulting in significant fatigue life enhancements, as demonstrated by representative coupon testing. A special manufacturing jig has been designed and transitioned to the RAAF, which has used it to repair housings. These housings have performed well in flight tests, with no cracking detected.

Pseudo Fatigue Testing for Rapid Certification to Service

Aircraft full-scale fatigue tests are expensive to conduct and they are a critical item on the certification path of any aircraft design or modification. Two aspects that contribute to the cost of a test are its duration and the loads spectrum development process. This paper provides a summary of a proposed supplemental pseudo full-scale fatigue test (FSFT) aimed at rapid certification. In this instance the method was developed with the aid of extant FSFTs that were found to be deficient. The proposed process involves the development of proof loads, damage size estimates, a loads application rig, insertion of the target damage or modifications and conducting proof testing. As all locations with a propensity to crack are known, the process is considered to be the equivalent of having conducted a representative fatigue test for the required service life target and then demonstrating adequate residual strength (i.e. proof testing the damage state at the end of a FSFT).