How would I do stress analysis on a fuselage like this?

Question

This is a frame for an RC plane that I have designed using balsa wood and prefabricated carbon fiber flat bars. It will have a balsa wood skin on the outside, but I have yet to figure out how to select the proper skin thickness using shear flow calculations. It is about 60" long.

I am trying to run stress calculations on it to ensure that it can withstand flight loads. I have read over some methods for calculating stresses in stringers and skins in a couple textbooks, but I am having difficulty using the methods on the one I have designed. Does anyone have any insight on where to start with something like this?

Also, are there any standard boundary conditions to use for fuselage analysis like this? I have been assuming that the fuselage is fixed at the wing and loads are applied to the frame of the fuselage where the internal cargo weights act.

Answer 1

Even without calculation, this:

...looks like a pretty serious weak area (and similarly on the front). You have a joint between materials with (almost certainly) considerably different stiffness. You're likely to see quite a bit more stress right at that joint.

If at all possible, I'd do something like tapering the end of the carbon fiber and wood spars where they meet, so you get a more gradual transition to avoid an abrupt change like that.