A composite laminate is not a material you look up — it is a structure you design, ply by ply. Two laminates of identical mass and fibre can behave completely differently.
Plies and orientation do the work
Fibres carry load along their length and very little across it, so orientation is the design. The convention: 0° along the primary load, ±45° for shear and torsion, 90° for transverse load and to stop splitting. A typical quasi-isotropic stack is [0/±45/90]ₛ, which behaves roughly the same in every in-plane direction — at the cost of never being optimal in any one.
Symmetry and balance — or it warps
Two rules prevent most manufacturing surprises. Symmetric: the stack mirrors about the mid-plane, which decouples bending from stretching. Balanced: every +45° ply has a matching −45°, which decouples shear from stretching. Break symmetry and the part warps as it cools out of the autoclave; break balance and it twists under load. This is why an asymmetric laminate is nearly always a mistake rather than a clever optimisation.
The weak directions
A laminate is strong in-plane and weak through the thickness — there are no fibres there, only resin. That makes interlaminar shear and peel the usual failure modes, and it is why composites hate out-of-plane loads, sharp ply drops and unsupported free edges. Never load a composite in through-thickness tension if you can design it out.
Practical rules
Keep ply drops gradual (one ply per ~20× thickness), avoid stacking more than ~4 identical orientations together, and support free edges. Design allowables come from CMH-17 (ex MIL-HDBK-17) and must be knocked down for environment — hot/wet conditions can cut matrix-dominated properties substantially. Barely-visible impact damage (BVID) is the usual design driver on aircraft structure, not static strength.