If you start with an off the shelf 6" by 6" by 3/8" thick wide flange pultruded
fiberglass beam that is 70% as heavy as aluminum and taper the flanges
with a saber saw you can make a pretty good spar in a few hours or so.

Tensile strength is up to 20,000 psi which is probably close
to aluminum on a strength to weight ratio basis.
The beauty of fiberglass is it is less likely to fatigue
and crack.

6 G's on a 2000 pound gross weight airplane with a wing
span of 25 feet would give a bending load of about
6000 foot pounds. Total span in inches is 300".
Take out 50" for the cockpit and we get 250".
each wing is therefore 125 inches with a load
of 6000 pounds or 48 pounds per inch distributed load.

First thing we need do is get the moment of inertia for
this beam. The formula is given in the attached png.
The moment of inertia is 97.

Now we need the stress so we use a cantilever beam calculator.
http://www.efunda.com/formulae/solid_mechanics/beams/casestudy_bc_cantilever.cfm
Use the distributed load model.

Tip deflection will be about 4 1/4  inch.
Max stress at 10,700 psi will be well under the claimed
max bending stress for this beam of 20,000 psi.
Six G factor of safety of 2.

Very carefully check my arithmetic and run this by another
mechanical engineer to double check everything. I could have
made a serious mistake. Also do a static bending check on the wing
before you fly it.

I suspect one could go down to 1/4 inch on the flange and
web thickness and run the numbers again.

Paul Lamar