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L Molent - One of the best experts on this subject based on the ideXlab platform.
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service fatigue cracking in an Aircraft bulkhead exposed to a corrosive environment
Engineering Failure Analysis, 2013Co-Authors: Simon Barter, L MolentAbstract:Abstract It has been postulated that for Combat Aircraft corrosion generally occurs whilst the Aircraft is on the ground, while fatigue damage in the form of cracking (if any) grows as result of operational loading which mainly occurs in flight where the conditions are cold and dry and the loading rates are high. This would suggest that (in general) for Combat Aircraft the effects of the environmental degradation and any fatigue crack growth are decoupled. From a damage prediction viewpoint, this will significantly simplify assessment of such problems (i.e. the growth of fatigue cracks in these Aircraft types is not environmentally assisted). This paper presents a supporting case study that examines fatigue cracks that were detected in an F/A-18 Hornet bulkhead during post-service testing and teardown. The in-service phase of the cracking had significant evidence of oxidation on their surfaces which indicated an exposure to a mildly corrosive environment. Both the service and laboratory phases of the cracking were the subject of quantitative fractography and estimates of the crack growth rates were made. A comparison of the in-service and the in-test phases of crack growth indicated that no notable effect on the service part of the fatigue crack could be attributed to its exposure to the service environment.
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calculating crack growth from small discontinuities in 7050 t7451 under Combat Aircraft spectra
International Journal of Fatigue, 2013Co-Authors: Rhys Jones, L Molent, S A BarterAbstract:Abstract This paper supplements previous work which showed that the crack growth rate in a large range of metallic materials fitted a variant of the Hartman–Schijve equation where da/dN is a function of (ΔK − ΔKthr)α, where ΔKthr is a parameter that reflects the apparent fatigue stress intensity threshold of the material, and α is approximately 2. For the case of 7050-T7451 aluminium alloy the same equation is shown to fit both long and short crack growth data once the appropriate ΔKthr is chosen for each specific data set. This equation is used here to produce accurate predictions of the fatigue crack growth in 7050-T7451 aluminium alloy specimens with both a low and high stress concentration subjected to two Combat Aircraft loading spectra. Thus, it is postulated that if long crack data are fitted to the variant of the Hartman–Schijve equation then accurate predictions can be made in the short crack regime for a wide range of materials.
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fatigue crack growth from flaws in Combat Aircraft
International Journal of Fatigue, 2010Co-Authors: L MolentAbstract:The 1969 in-flight failure of an F-111 wing due to the presence of a rogue manufacturing flaw precipitated the adoption of the damage tolerance Aircraft design philosophy by the United States Air Force. This article compares this flaw to several flaws or initiating discontinuities that led to unexpected failures in Combat-type Aircraft operated by the RAAF and found these to be in general an order of magnitude smaller. Failure in the RAAF examples considered here were also dependent upon several other design and production factors in-addition to the initial discontinuity size.
R Sunder - One of the best experts on this subject based on the ideXlab platform.
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engineering analysis of notch root fatigue crack growth under spectrum loading
International Journal of Fatigue, 1991Co-Authors: R SunderAbstract:An engineering procedure is described to predict notch root fatigue crack growth under spectrum loading. Long-crack growth rates and closure data, notch root stress-strain analysis and Rainflow cycle counting constitute major inputs to the prediction process. The procedure models non-uniform extension of the part-through crack front, crack growth in compression fatigue and the effect of tensile and compressive overloads. Crack growth predictions under constant-amplitude and Combat Aircraft spectrum loadings were compared with data from the literature for 2.3 mm and 6.25 mm thick 2024-T3 alloy sheet material. Good predictions were obtained for the thicker material using the linear elastic fracture mechanics approach. The thinner material required inclusion of strain-based intensity and residual stress effects for more realistic predictions. Neglecting these factors as a rule provided life estimates that are not conservative.
S Pradeep - One of the best experts on this subject based on the ideXlab platform.
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nonlinear control of unmanned Combat Aircraft during take off
40th AIAA Aerospace Sciences Meeting & Exhibit, 2002Co-Authors: S PradeepAbstract:Unmanned Combat Aircraft have been projected to provide significant improvements in performance over their manned counterparts, without endangering the crew, thus offering a cheaper and more viable option. A major challenge in this nascent area is the design of the flight control system. This paper deals with the design of a feedback linearized controller for automating take-off, which is an inherently nonlinear phase of flight. Pilot opinion was codified to serve as a standard for control design. Based on this, the Aircraft was made to track a predetermined pitch rate, accounting for saturations in the the elevator and elevator rates. Simulations show that the performance of the controller is acceptable in the presence of plant uncertainty and sensor noise. Of the plant parameters, variations in the centre of gravity affect it most. Response in the presence of sensor noise is good. A Monte-Carlo simulation shows that the controller does not exhibit unstable tendencies and its performance is acceptable. With these promising results, feedback linearization is proposed as a good nonlinear control design tool for automating Combat Aircraft take-off. © 1998 by S. Pradeep.
Rhys Jones - One of the best experts on this subject based on the ideXlab platform.
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the usaf characteristic k approach for cracks growing from small material discontinuities under Combat Aircraft and civil Aircraft load spectra
Engineering Failure Analysis, 2017Co-Authors: Pu Huang, Daren Peng, Rhys JonesAbstract:Abstract This paper investigates the ability of the USAF characteristic K approach to represent the growth of cracks from small naturally occurring material discontinuities. The examples discussed in this paper are associated with two different load spectra. The first case studied involves crack growth in AA7050-T7451 under a Combat Aircraft load spectrum. The second case studied the crack growth in AA7050-T7451 under a civil Aircraft load spectrum. It is shown that using the root mean square value of ΔK in a load block to characterize the stress intensity factor together with the NASGRO equation, with the constants determined from constant amplitude tests on long cracks, enables the growth of small cracks under these two different complex load spectra to be reasonably accurately represented.
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calculating crack growth from small discontinuities in 7050 t7451 under Combat Aircraft spectra
International Journal of Fatigue, 2013Co-Authors: Rhys Jones, L Molent, S A BarterAbstract:Abstract This paper supplements previous work which showed that the crack growth rate in a large range of metallic materials fitted a variant of the Hartman–Schijve equation where da/dN is a function of (ΔK − ΔKthr)α, where ΔKthr is a parameter that reflects the apparent fatigue stress intensity threshold of the material, and α is approximately 2. For the case of 7050-T7451 aluminium alloy the same equation is shown to fit both long and short crack growth data once the appropriate ΔKthr is chosen for each specific data set. This equation is used here to produce accurate predictions of the fatigue crack growth in 7050-T7451 aluminium alloy specimens with both a low and high stress concentration subjected to two Combat Aircraft loading spectra. Thus, it is postulated that if long crack data are fitted to the variant of the Hartman–Schijve equation then accurate predictions can be made in the short crack regime for a wide range of materials.
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implications of the lead crack philosophy and the role of short cracks in Combat Aircraft
Engineering Failure Analysis, 2013Co-Authors: Rhys Jones, Dinaz Zenobia TamboliAbstract:Abstract The Australian Defence Science and Technology (DSTO)/Royal Australian Air Force (RAAF) approach to the management of fatigue cracking in Combat and trainer Aircraft makes use of the “lead crack” concept. In this approach crack growth is assumed to initiate from small naturally occurring defects/discontinuities with dimensions of approximately 10 μm and growth is assumed to commence from day one. As a result, for certification purposes, we need to address the so called short crack anomaly, whereby for a given Δ K the crack growth rate ( da / dN ) is significantly greater for short cracks than it is for long cracks. In this paper we reveal that there are several instances where this anomaly vanishes if da / dN is represented as a function of (Δ K − Δ K thr ), where Δ K thr can be thought of as an apparent threshold. We then show that for operational Aircraft the “true” da / dN versus Δ K curve is an amalgam of the short and long crack growth curves. We next show that existing test procedures used to establish the effect of composite repairs to cracks in fleet Aircraft overestimate their effect. We also show that the growth of cracks from small naturally occurring defects exhibits little, if any, R ratio effects.
V R Ranganath - One of the best experts on this subject based on the ideXlab platform.
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spectrum modification for full scale fatigue testing of an ageing Aircraft
2006Co-Authors: C M Manjunatha, V R RanganathAbstract:The flight load sequence of a Combat Aircraft was modified by eliminating different levels of small amplitude load excursions to derive several different test load sequences. The fatigue crack growth behavior under all these spectrum load sequence was predicted in a single edge notched tension specimen of an airframe grade La16 aluminum alloy. Crack growth behavior was predicted using a fatigue crack growth law derived from the constant amplitude fatigue crack growth tests, incorporating crack closure effects. It was observed that full scale fatigue testing time can be reduced significantly by using of one of the derived test load spectrum without compromising, in general, on the fatigue damage caused by eliminated load cycles.
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effect of airframe materials on spectrum modification for full scale fatigue testing sf an ageing Aircraft
2006Co-Authors: C M Manjunatha, V R RanganathAbstract:In order to accelerate full scale fatigue testing, the original flight load spectrum of a Combat Aircraft was modified by eliminating small amplitude load excursions lower than a prefixed filter value. Different levels of filter value were used to derive several different test load spectrum having lesser number of fatigue cycles than the flight load spectrum. The fatigue crack growth behavior under all these spectrum load sequences was predicted in two different airframe materials viz., D16 aluminum and Ti6A14V alloys using a fatigue crack growth model derived from constant amplitude fatigue crack growth tests, which incorporated crack closure effects. Results show that the basis of spectrum modification should not only be the elimination of small amplitude load cycles but also on the fatigue crack growth behavior of the structural materials used in airframe construction.