Composites

Generic Composite Panel / Stringer Analysis

Sample Calculations are shown here as an example to demonstrate, the software capability to calculate any combination of number of ply layers, thickness, angle, material and load combinations. (material data shown here does not represent any specific material !)

The software can be used for sizing of aircraft composite laminate panels, wing skins, composite stringers and fuselage panels. The software is very user friedly and only requires material data, loads and ply orientation as input. It will calculate the critical ply failure, the margins and the mode of failure. It can also be used for residual strength estimation for damage tolerance analysis. Please do not interpolate or extrapolate these sample calculation for any design purpose.

The application can be used to size composite wing stringers and skin for different stringer shapes (I, J and omega shape). The application is developed to minimize B Matrix to zero and D16 and D26 ( in the ABD Matrix ) to zero for preventing warping during curing. User inputs the loads and the desired thickness and the applications develops the optimum ply lay up and critical margin of safety.

Please click on the image below for video # 1 to begin.

Please click on the image below for video # 2 to begin.

Sample Screen prints from Composite Application.

Composite Skin Panel Output ( Strain Output for each layer and Margins )

Composite Skin Results Summary (with ABD Matrix)

Composite Panel Buckling