Select composite material type (carbon/epoxy, glass/polyester, aramid blend) from the material library
Input ply orientation angles (0°, 45°, 90°) and stack sequence using standard notation (e.g., [0/45/-45/90]s for symmetric laminate)
Define applied loads (in-plane tension, shear, bending moment in kN or lbf) and boundary conditions
Run the classical laminate theory (CLT) solver to generate stiffness matrices and stress/strain distributions
Review failure criterion results using Tsai-Wu, Hashin, or Puck criteria against material allowables

Worked Example

Carbon/epoxy quasi-isotropic laminate [0/±45/90]s with ply thickness 0.125 mm, total laminate thickness 1.0 mm. Applied tensile load 50 kN/m width. Using CLT with E1=140 GPa, E2=10 GPa, G12=7 GPa, ν12=0.3 per ply: mid-plane strain εx=0.00368, max ply stress in 0° plies=515 MPa. Hashin matrix failure criterion predicts failure initiation at 62 kN/m; Tsai-Wu reduces margin to 58 kN/m accounting for biaxial effects. Reserve factor = 1.16 for this load case.

Practical Notes

Angle-ply contributions (±45°) dominate shear stiffness; symmetric/balanced laminates minimize warping and thermal distortion in aerospace and wind turbine blade applications
Thin laminate assumption (thickness << in-plane dimensions) breaks down below 0.5 mm; apply 3D FEA for edge delamination validation near free boundaries
Environmental degradation (moisture uptake in epoxy absorbs ~3% weight over months in humid climates) reduces matrix properties; adjust E2 and G12 downward by 15–25% for knockdown factors
Ply drop zones in tapered sections create stress concentrations; CLT predicts average strains only—use local 3D analysis for microbuckling onset in compression-dominated regions