Download 1) It is required to provide a life estimate for a wing lower skin joint

1) It is required to provide a life estimate for a wing lower skin joint using the S-N curve provided
in Sketch 6.1 (Figure 1 (a)). The cumulative frequency versus bending moment curves for the
aircraft were determined in the examples given in ESDU Data Items No 69023 and 75008 and are
presented here together in Sketch 6.2 (Figure 1(b)). At the particular skin joint under consideration
the section modulus is 0.0666 m3
(a)
(b)
Figure 1
2)Two cylindrical pressure vessels made of the same unidirectional lamina but with different fiber
orientations were loaded as shown and gave the following strain readings (Figure 2):
Figure 2
Assuming   0.3 , determine lamina module E1 , E2 , G12 .
3) Show that [0/±60]s laminate is quasi-isotropic, i.e., Axx  Axx for any angle  (Figure 3).
Figure 3
4) A [±30]s laminate is loaded in uniaxial compression as shown (Figure 4). Determine the
compressive strength Fxc at FPF (first ply failure) according to the maximum strain theory for the
following given properties:
1ut  0.015, 1uc  0.015,  2ut  0.006,  2uc  0.024,  6u  0.015, Ex  61.4GPa,  xy  1.2
Figure 4
5) Determine exact and approximate values for the strength Fxt of the above problem (4) for the
carbon/epoxy material AS4/3501-6 listed below using the maximum stress theory. What is the
prevailing failure mode? The laminate modulus is Ex  64.1GPa and Poisson´s ratio is  xy  0.65
.
Table 1
6) Determine the stress concentration factor for the joint schema showed in Figure 5. See Michael
Niu chapter 7. Fastner HL10VK-8 ( 1 " ), width w=50,0 mm, thickness t=10,0mm.
4
Figure 5
7) At a location of interest in a member made of the Ti-6A1-4V alloy of Figure 6 (b), the material is
repeatedly subjected to the uniaxial stress history shown in Figure 6 (a). Estimate the number of
repetitions necessary to cause fatigue failure.
(a)
(b)
Figure 6