The engineering for aluminum has numerous texts with worked examples (actual calculation of sub-assemblies). **Maybe I need to see a similar texts for composites. **I believe that much of the data is for autoclaved parts, which is usually more than a homebuilder does. The composites are always being improved and there are more resins every day with older ones being retired. Aircraft Aluminum and steel has not changed much since the 40's and much of the data on it is actually from that period. I also marvel at the CF stuff but having worked the "top of the line" with the AV8-B (Harrier) I also know those results would be much more difficult to achieve without the "pre-pregs" and 250 PSI high temperature autoclave.

The math is the same. The solutions are different.

Take a spar. Calculate I and stresses in the caps. Calculate local shear. Check web and caps for buckling.

A CFRP spar is exactly the same math, except that buckling for the web (and often) the caps is avoided by simply making the core thick enough.

Pre-pregs are certainly stronger, but not much stiffer, so their overall weight reduction isn't nearly as much as one would expect at first sight. The LBA allowables are achievable by any decent "amateur" and have 35 years of proven history, so I wouldn't sweat those too much if you're building technique is solid.

If we get away from "black aluminium", it gets simpler. Take a monocoque, sparless wing. Simply calculate local bending and shear in all 3 directions, check stresses and do a single buckling check on the part of the skin with the highest stresses. Thicken the core there enough to not buckle and you're done with the calculations.

Here's an example that covers about a third of the math for a full wing. You obviously have to run it through multiple load cases (flaps up/down, ailerons up/down, gust, combined loads etc, which means that local lift will vary non-linearly), but that's simple iteration.