Electronic enclosures for space or avionics application can be designed usi
ng laminated composites to reduce weight, provide a modular design that has
equal or better thermal and mechanical performance, and has a lower cost p
er enclosure than the standard "black'' aluminum design. Initial sizing of
an enclosure to determine the number of plies and ply orientation can be ac
complished by subdividing the structure into simple shapes and analytic clo
sed-form equations used to calculate bending stresses and deflections, unia
xial and shear buckling allowables, and natural frequencies. This initial s
izing was performed on a three-sided enclosure with integral mounting flang
es. The walls were analyzed using static equivalent of random vibration loa
ds in closed-form analytic approximate or exact equations and compared with
those using finite element analysis (FEA). Depending on the degree of orth
otropy, i.e., how close the off-diagonal flexural stiffnesses are to zero,
the analytic predictions for laminae stresses vary with finite element resu
lts. Two different hybrid PAN/pitch fiber/epoxy laminates and a carbon fabr
ic/epoxy laminate with varying degrees of orthotropy were chosen for compar
ison. The margins of safety for the analytic results was within 5% of the F
EA results for the orthotropic laminate but was different by factors of 1.5
to 13 for the non-orthotropic laminates. There was good comparison between
analytic solutions and FEA for buckling, natural frequency, deflection, an
d stresses, in all cases the analytic predictions were conservative. These
analytic equations were used for initial sizing of an enclosure, and a deta
iled FEA was performed on the electronics enclosure under actual random vib
ration loads. The final enclosure was fabricated and tested under these ran
dom vibration loads. (C) 1999 Elsevier Science Ltd. All rights reserved.