Page 448 - Aircraft Stuctures for Engineering Student

10.6 Laminated composite structures 429
From Eq. (10.48) the modulus of the bar is given by

80 x 10 80 x 40
+ 5000 x -
E1 = 200000 x -
80 x 50 80 x 50
i.e.
El = 44000N/m111'

The direct stress, q, in the longitudinal direction is given by
100 x io3
a1 = = 25.0 N/m2
80 x 50
Therefore, from Eq. (10.45), the longitudinal strain in the bar is
25.0
5.68
&I = - x 10-~
=
44 000
The lengthening, A,, of the bar is then
AI = 5.68 x x 500
i.e.

A1 = 0.284-
The major Poisson's ratio for the bar is found from Eq. (10.50). Thus
80 x 40 x 0.2 + -
'80 x 10
x 0.3 = 0.22
vlt =m 80 x 50
Hence the strain in the bar across its thickness is

&, = -0.22 5.68 10-~ = -1.25 10-~
The reduction in thickness, A,, of the bar is then

A, = 1.25 x x 50
i.e.
A, = 0.006111m
The stresses in the epoxy and the carbon are found using Eqs (10.46). Thus
nrn (epoxy) = 5000 x 5.68 x = 2.84N/mm2

of (carbon) = 200000 x 5.68 x = 113.6N/mm2

10.6.2 Composite plates

In Chapter 5 we considered thin plates subjected to a variety of loading conditions.
We shall now extend the analysis to a lamina comprising a filament and matrix of
the type shown in Fig. 10.66.