Page 395 - Aircraft Stuctures for Engineering Student

376 Stress analysis of aircraft components

Table 10.2

1 381.0 302.4
2, 16 352.0 279.4
3, 15 269.5 213.9
4, 14 145.8 115.7
5, 13 0 0
6, 12 -145.8 -115.7
7, 11 -269.5 -213.9
8, 10 -352.0 -279.4
9 -381.0 -302.4

For this particular section Zxy = 0 since Cx (and Cy) is an axis of symmetry.
Further, My = 0 so that Eq. (9.6) reduces to

in which
Zxx = 2 x 216.6 x 381.02 + 4 x 216.6 x 352.02 + 4 x 216.6 x 26952

+ 4 x 216.7 x 145.82 = 2.52 x lo8 mm4
The solution is completed in Table 10.2.

10.2.2 Shear

For a fuselage having a cross-section of the type shown in Fig. lO.ll(a), the
determination of the shear flow distribution in the skin produced by shear is basically
the analysis of an idealized single cell closed section beam. The shear flow distribution
is therefore given by Eq. (9.80) in which the direct stress carrying capacity of the skin
is assumed to be zero, i.e. tD = 0, thus

Equation (10.17) is applicable to loading cases in which the shear loads are not
applied through the section shear centre so that the effects of shear and torsion are
included simultaneously. Alternatively, if the position of the shear centre is known,
the loading system may be replaced by shear loads acting through the shear centre
together with a pure torque, and the corresponding shear flow distributions may be
calculated separately and then superimposed to obtain the final distribution.

Example 10.5
The fuselage of Example 10.4 is subjected to a vertical shear load of 100 kN applied at
a distance of 150 mm from the vertical axis of symmetry as shown, for the idealized
section, in Fig. 10.12. Calculate the distribution of shear flow in the section.