calculation
I pulled an equation out of the machinists holy bible... A big, old, green book we have in the shop on campus. It gave cross sectional information on many, many different shapes. I picked out a formula for a symmetric "C" channel, even though the top spar bar is about .125" shorter than the bottom. I used the top spar dimension as this would give a conservative (smaller) figure.
I=[(B*D^3)-H^3(B-T)]/12
B=.125+.063+1.25=1.438"
D=7.75"
H=7.75-(1.375)*2=5"
T=.125+.063=.188"
SUBSTITUTED IN: I = 42.75"^4
I assumed the skin to provide no additional support to the spar. This was done so that any additional strength from the skin would be left in the reserve margin of the structure (fatigue, corrosion, overloading, poor construction, etc.)
BENDING LOADS: Q=(M*Y)/I ALSO, Mmax=(Qmax*I)/Y
Q=yield strength=60ksi (2024-T3)
M=bending moment
Y=distance from neutral axis to extreme fiber
I=moment of inertia=42.75"^4
Mmax = [(60,000psi*42.75)/3.875"] = 661935 in*lb = 55160 ft*lb
So the wing has about a 10' arm, and if I assume a constant uniform lift distribution... (worse than real world bending loads?)
M=F*D and F=M/D 55160 FT*LB/5' = 11032 LB
SPAN LOADING: 11032#/10' = 1103.2# PER FT
WING LOADING: SPAN LOADING/CHORD = 1103.2/4.8' = 229.8#/FT^2
Now for the good stuff that makes a little more sense...
Wing loading * wing area = loading so, 229.8#/ft^2*120ft^2 = 27580#
and loading/gross weight = g forces so, 27580#/1800# = 15.3gs
Any comments? I'm in no way a structure genious and there have to be many things i've left out. I hope anyone interested could give there input here and maybe we could refine this number even further... Heck, I might even be so far off none of this is even worth typing!!! I just don't know.
Hope to hear more thoughts here,
Jonathan
RV-7A finishing Fuselage