PSI - Issue 17

T. Martins et al. / Procedia Structural Integrity 17 (2019) 878–885 8 Martins, T., Infante, V., Sousa, L., Antunes, P.J., Moura, A.M., Serrano, B./ Structural Integrity Procedia 00 (2019) 000 – 000 size. The factors chosen by the manufacturer will be kept in these proposals, making the suggested time for first inspection 25896/3 = 8632 FH. The manufacturer defined a crack of 1.5mm in the frame to be of critical size, which cannot be exceeded in the aircraft's operating life. For a crack size of 30mm, the part is in the imminence of failure. The manufacturer suggests the use of a safety factor of 6 to the time a crack of 1.5mm would take to grow to these 30mm. From the adjusted curve, a crack size of 1:5mm is reached for 45525 FH and after 61901 FH a crack of 30mm is expected to be found. From these, the suggested time between subsequent inspections is 2729 FH. 5. Conclusions The author proposes that first inspection for crack detection in the frame is performed after 8500 FH, with the following inspections executed in intervals of 2700 FH each, until a crack size of 1:5mm is detected, signalling the end of the aircraft's useful life. These are a significant reduction from the manufacturer's proposed periods but should prevent premature failure due to the more severe PoAF load spectra. From the methods used, the authors conclude that: • The use of XFEM with a reduced mesh overestimates the SIF results obtained by FEM. A more stable solution for the extended method would require a finer mesh than the 0:5mm element size used around the crack tip ( 1 ≤ ≤ 6 ) • For the propagation laws, NASGRO tends to estimate the longest life because of the crack opening function. Paris law estimates the shortest lifetime. Test specimens for negative stress ratios may be necessary for results from the cycle-by-cycle integration closer to those obtained in the CEAT tests [1]. 6. Acknowledgements This work was supported by “Fundação para a Ciência e a Tecnologia” (FCT), through the Institute of Mechanical Engineering (IDMEC) under the Associated Laboratory for Energy, Transports and Aeronautics (LAETA), Project UID/EMS/50022/2013. [1] CEAT. Epsilon - documents de synthese suite a la campagne d'essai de fatigue. Technical report, SOCATA, Groupe Aeroespatiale, 1996. [2] B. A. S. Serrano, V. Infante, and B. S. D. Marado. Fatigue life time prediction of poaf epsilon tb-30 aircraft - part i: Implementation of different cycle counting methods to predict the accumulated damage. In Iberian Conference on Fracture and Structural Integrity - CIFIE, 2010. [3] Milharadas. Relatório de tirocínio. Technical report, Força Aérea Portuguesa, 2004. [4] V. de Brederode. Fundamentos de aerodinamica incompressivel, chapter 10, pages 563-567. IST Press, 2014. [5] R. J. Hartranft and G. C. Sih. An approximate three-dimensional theory of plates with application to crack problems. International Journal of Engineering Science, 8, 1970. [6] I. S. Raju and J. C. Newman. Stress-intensity factors for two symmetric corner cracks. In Fracture Mechanics; Proceedings of the Eleventh National Symposium, 1979. [7] B. A. S. Serrano, V. Infante, and B. S. D. Marado. Fatigue life time prediction of poaf epsilon tb-30 aircraft - implementation of automatic crack growth based on 3d finite element method. Engineering Failure Analysis, 33:17 - 28, 2013. [8] A. C. Pickard. The application of 3-dimensional finite element methods to fracture mechanics and fatigue life prediction. Warley,U.K., 1986. [9] D. Broek. Elementary Engineering Fracture Mechanics. Sijthoff & Noordhoff International Publishers B.V., 1978. [10] N. E. Dowmling. Mechanical Behaviour of Materials. Pearson Education Limited, 2013. [11] Society of Automotive Engineers. SAE Fatigue Design Handbook, 1st edition, 1968 [12] S. R. Mettu, J. M. Shivakumar, F. Beek, L. C. Williams, R. G. Forman, M. J. J., and J. C. Newman. Nasgro 3.0: A software for analyzing aging aircraft. In The Second Joint NASA/FAA/DoD Conference on Aging Aircraft. 885 References