L. Molent

On the Glinka and Neuber methods for calculating notch tip strains under cyclic load spectra

Abstract To maintain the continued airworthiness of civil and military aircraft it is essential that the fatigue behaviour of components subjected to complex multiaxial stress conditions be both understood and predicted. This topic is extremely complex and numerous criteria ranging from the purely empirical to the theoretical have been proposed. To this end it is necessary to know the localised stress–strain history. Whilst a detailed finite element analysis can be performed it is often very time consuming. Consequently, the present paper focuses on the development of a simple method which combines modern constitutive theory with either Neubers, or Glinkas, approach to calculate the localised notch strains. Here particular attention is paid to the effects of cyclic loading, mean stress relaxation and its effect of fatigue damage.

Design of an all boron/epoxy doubler reinforcement for the F-111C wing pivot fitting: Structural aspects

Abstract This paper presents an overview of the structural aspects of the design and development of a local reinforcement designed to lower the stresses in a region of the F-111C wing pivot fitting which is prone to cracking. The stress analysis, representative specimen testing, thermal analysis and aspects of the full-scale static testing of this design are summarized.

Study of multi-site damage of fuselage lap joints

The downturn in the world economy coupled with the cost of new aircrafts has meant that there are now ageing fleets whose continued airworthiness requires special attention to corrosion treatments, repair design, fatigue and fracture analysis, and improved crack detection techniques. To assist in this goal the present paper first summarises recent Australian efforts into the development of a simple experimental test specimen which is capable of reproducing the crack growth and failure mechanisms seen in the fuselage lap splice of a wide bodied transport aircraft. The development of a composite repair to overcome these phenomena is then discussed.

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.

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.

Repair of Multi-Site Damage

Publisher Summary This chapter presents the results of a fatigue test program, which also considers environmental and damage-tolerance aspects, conducted using specimens representative of wide-bodied commercial aircraft fuselage lap joints. The phenomenon of multi-site damage (MSD) in aircraft has been under examination by many in the aviation industry. This chapter investigates the feasibility of applying advanced bonding technology to commercial aviation structures containing MSD. The consequence of the undetected presence of MSD was dramatically illustrated through the in-flight failure of a fuselage lap joint on an Aloha Airlines B-737 aircraft on April 28, 1988. Essentially this failure occurred due to numerous small cracks along a fastener line linking together, causing the residual strength of the fuselage to be exceeded under pressurization. A test program was conducted to reproduce this type of failure, and an adhesively bonded boron/epoxy doubler for use as a repair or preventative measure has been developed.

The F111C Wing Pivot Fitting Repair and its Implications for the Design/Assessment of Bonded Joints and Composite Repairs

Abstract This chapter discusses the development of a bonded composite reinforcement to the integrally stiffened D6ac steel wing pivot fitting (WPF) of F-111 aircraft while in-service with the Royal Australian Air Force (RAAF). Particular attention is given to the building block approach and as such describes the finite element and experimental program used to determine improved stiffener runout geometry as well as the full-scale wing tests required to validate the computed stress reduction. Other aspects of this reinforcement are discussed in this chapter. This represents the first bonded reinforcement to primary structure and was successful in that by reducing the strains at the critical location by approximately 30% it achieved its primary objective namely to stop cracking and wing failure in the RAAF F-111 fleet and Cold Proof Load Test (CPLT; explained later). A significant outcome of the F-111 WPF reinforcement program was that it resulted in a detailed understanding of the design/assessment tools needed for designing damage tolerant repairs to meet the stringent damage tolerant requirements inherent in both JSSG2006 and FAA ac-120-107b. It is also shown that the in-service performance of the WPF reinforcement reveals shortcomings in the approach outlined in the US Composite Materials Handbook CMH-17-3G for the design of bonded joints. It is also shown that not only was the WPF reinforcement the first example of the ability of composite repair technology to eliminate cracking in primary structure it was also the first example to highlight the benefit to be gained by combining composite repair technology with shape optimisation or rework.

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).

Determination of material constants for crack size-dependence crack growth model

A crack size-dependence crack growth model was used to characterise the fatigue crack growth in AL 7075-T6 and AL 2024-T351 alloys. It is shown that the crack growth parameters %K and a can be used to linearise the crack growth in regions I (elastic) and II (plastic) by plotting fatigue data linear-linear da/dN×√a versus %K3, where cubic stress dependency is assumed. A theoretical attempt was made to relate this crack size-dependence fatigue crack growth parameters to the strain-life relationship constants. A reasonably good agreement was achieved when comparing between the theoretical predicted and experimental determined material constants.