PSI - Issue 36

M. Karuskevich et al. / Procedia Structural Integrity 36 (2022) 92–99 M. Karuskevich, T. Maslak, Ie. Gavrylov et al. / Structural Integrity Procedia 00 (2021) 000 – 000

96

5

5.3 kN, which is approximately 75% of 5.8 class bolts’ yield stress. Frictional contacts between parts are introduced with appropriate friction coefficients – 0.3 for steel-steel contact and 0.61 for aluminium-steel contact.

Fig. 4. Simplified model and boundary conditions of the indicator attached to the aircraft structural component (specimen for fatigue test).

Calculated displacement along the y direction is presented in Fig. 5. It can be seen that due to the bolted connection, the deformation of the specimen follows the deformation of the element.

Fig. 5. Displacement along y direction.

Fig. 6 shows the minimum principal stress. One can see the compression of the material surrounding the holes. It is a result of bolts’ preload transferred by washers. Compression is higher on the holes’ edges along the y direction due to pressure caused by bolts. There is no compressive stress in the specimen’s work section and the element. In Fig. 7 equivalent stress distribution in the fatigue indicator is shown. There is no significant stress concentration in the neighbouring of holes, and maximum stress occurs in the “neck”. Whilst the task for sensitivity increase looks realistic by manipulation with geometry, the problem of buckling prevention introduces limitations on the indicator practical use. The integral characteristic of the accumulated fatigue damage is provided by the wing fatigue assessment. The problem is that the wing components work both under the compression and tension. In flight the lower panels of wing structure carry tension loads, whilst on the ground, these components work under the compression, and vice versa, upper panels work under compression in air, and under the tension being on the ground. This situation causes limitations on the fatigue indicator geometry variation and requires additional buckling analysis. This analysis has been conducted by two methods: a) following the classic Euler’s formula for critical load; b) using eigenvalue buckling analysis in ANSYS software. Limitations on the compression loads have been calculated according with the FEM. It was found that the danger of the stability loss exists only for the compression stress in the structural component equal to 140 MPa. Taking into account the realistic spectrum of loads on aircraft wing this value is beyond the operational level and can be ignored. Possibility for the sensitivity optimization of the fatigue indicator is illustrated by stress distribution of two