Issue 35

R. Sepe et alii, Frattura ed Integrità Strutturale, 35 (2015) 534-550; DOI: 10.3221/IGF-ESIS.35.59 546 To simulate the effect of the breakage of the frame, another FE model of the panel has been developed. In this model the frame in the failure zone was not modeled. In Fig. 16 the panel zone is shown with the broken frame. In Tab. 6 some values of measured strains corresponding to a load applied along the frame direction ( P y = 400 kN) and the corresponding numerical ones, expressed with the relative deviation, are reported. The numerical results are in good agreement with the experimental results, the higher deviation of the strain gauge S3 is due to the secondary bending. Strain gages Uniaxial load along frame direction (T1 test) P y = 400 kN FEM strain (a) [μm/m] Experimental strain (b) [μm/m] Deviation = (a-b)*100/b [%] S1 -389 -460 7.7 S2 1312 1399 -6.2 S3 848 1305 -35.0 S4 1697 1727 1.7 S5 - - - S6 1317 1285 2.6 S7 -406 -420 3.2 Table 6 : Numerical and experimental correlation for load configuration T1 for panel with broken frame. Fatigue testing The results of conducted fatigue analysis of the fuselage panel are presented in form of contour plot of the ratio:                                 , , ( ) 1, 2, 3, , , 1 / f a eq a vonMises m m m f a p a (6) accordingly to the investigated criteria. The values equal or greater than 1 are indicative of critical points; the higher the value, the higher the possibility of failure. Values less than 1 are typical of locations that are not critical fatigue-wise. The resulting contour plots presents a consistent pattern, showing that the highest Sines ratio is always located at the bottom of the panel (marked by a circle in Figs. 17-21), near the third frame, on the rear side. As shown in Fig. 17, there are no critical points after 10˙000 cycles ( r max = 0.8); critical zones begin to appear just before 100˙000 cycles ( r max = 1.015) (Fig. 18). The extension of the critical zones increases at the locations near the lower bay, extending along the lower stringer. As the number of cycles goes up, more zones interested by the fatigue critical condition appear on the left side of the panel (Figs. 19-21). At 177˙000 cycles (Fig. 19) numerical results predict panel failure at an applied load level 1.033 times the experimental one. Figure 17 : ρ values after 10 4 cycles: front (left) and rear (right) sides.