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Simulation of Fatigue Crack Growth in Integrally Stiffened Panels Under the Constant Amplitude and Spectrum Loading

Augustin P.

Fatigue of Aircraft Structures

2009

Iss. 1

5--19

The paper describes methodology of numerical simulation of fatigue crack growth and its application on integrally stiffened panels made of 2024-T351 aluminium alloy using high speed cutting technique. Presented approach for crack growth simulation starts by the calculation of stress intensity factor function from finite element results obtained using MSC. Patran/Nastran. Subsequent crack growth analysis is done in NASGRO and uses description of crack growth rates either by the Forman-Newman-de Koning relationship or by the table lookup form. Three crack growth models were applied for spectrum loading: non-interaction, Willenborg and Strip Yield model. Relatively large experimental program comprising both the constant amplitude and spectrum tests on integral panels and CCT specimens was undertaken at the Institute of Aerospace Engineering laboratory in order to acquire crack growth rate data and enable verification of simulations. First analyses and verification tests of panels were performed under the constant amplitude loading. For predictions of crack growth using the spectrum loading a load sequence representing service loading of the transport airplane wing was prepared. Applied load spectrum was measured on B737 airplane within the joint FAA/NASA collection program. The load sequence is composed of 10 flight types with different severity analogous to the standardized load sequence TWIST. Before application on the stiffened panels a calculation of crack growth under the spectrum loading was performed for simple CCT specimen geometry. The paper finally presents comparison of simulations of fatigue crack propagation in two-stringer stiffened panel under the spectrum loading with verification test carried out in the IAE lab. The work was performed within the scope of the 6th Framework Programme project DaToN - Innovative Fatigue and Damage Tolerance Methods for the Application of New Structural Concepts.

Damage tolerance evaluation of the wing spar with cracked web

Augustin P.

Research Bulletin / Warsaw University of Technology, Institute of Aeronautics and Applied Mechanics

2001

nr 11

7-12

Damage tolerance evaluation is required for the aircraft life determination according to FAR Part 25 regulations. This concept is based on the crack propagation and residual strength determination. The damage tolerance evaluation of an aircraft structure, results in the inspection program ensuring that fatigue initiated cracks would involve no failure prior to the detection. The inspection program must be developed for each main structural element in the aircraft. Solutions to this problem usually found in the literature deal mainly with the components subject to tension; such as stiffened panels or lugs. This paper presents the testing program and FEM analysis focused on the wing spar with cracks in the web. It is characteristic of fatigue cracks in the web that they usually change their growth direction due to a complex stress state. The crack shape measured was used for constructing of a FEM model of the spar, the residual strength was calculated and compared with the experimental results.