Detail work on structures analysis

Among a bunch of great questions/suggestions/feedback, I took a look at the fore/aft distribution of forces on the two spar tubes.  The previous analysis ran a simple 50/50 split of the wing total forces to get things started.  What is better?  AVL can help provide the answer.

Under a 7g (using this now instead of 6g) load case at CL of 1.5 (assumed CLmax), the element forces can be plotted (seen above).  This gives the distribution of CP along the airfoil chordwise and showing the distribution of force.  For the root, here is the element force distribution:

 Strip # 25     # Chordwise = 10   First Vortex =241
    Xle =   3.16670    Ave. Chord   =    5.0000   Incidence  =    1.7500 deg
    Yle =  -0.59947    Strip Width  =   1.20000   Strip Area =    5.999976
    Zle =   4.03135    Strip Dihed. =   -2.9939

    cl  =   1.56703       cd  =   0.73643      cdv =   0.67992
    cn  =   1.66386       ca  =   0.47903      cnc =   7.78287    wake dnwsh =   0.10109
    cmLE=  -0.55324    cm c/4 =  -0.13727

    I        X           Y           Z           DX        Slope        dCp
  241     3.19462    -0.59918     4.03134     0.17282     0.28424     5.05604
  242     3.41428    -0.59918     4.03134     0.37261     0.14534     2.82108
  243     3.83407    -0.59918     4.03134     0.54628     0.08312     2.09844
  244     4.41670    -0.59918     4.03134     0.67141     0.03037     1.77600
  245     5.11040    -0.59918     4.03134     0.73689    -0.01091     1.54034
  246     5.85353    -0.59918     4.03134     0.73689    -0.04508     1.32883
  247     6.58005    -0.59918     4.03134     0.67141    -0.06874     1.10233
  248     7.22542    -0.59918     4.03134     0.54628    -0.08517     0.86729
  249     7.73230    -0.59918     4.03134     0.37261    -0.09597     0.62162
  250     8.05563    -0.59918     4.03134     0.17282    -0.10036     0.35427

The X position does not start at zero because I use the nose of the glider as the origin location.  An X position of 3.1667 is the LE at the root if you were curious.  Plotting the X position vs the dCp value gives the distribution of pressure along the airfoil chord.

Now some simple statics can give the force distribution between the LE and TE.  Assuming each CP acts like a force on a beam, the component of load carried by the LE and TE can be found by multiplying the force by a ratio of distance to the beam length.  The beam looks like this:

Component of force F acting at a distance xa from simple support A is given by: F_a = F * xb / (xa + xb).  Carrying out this algebra using the CP distribution from AVL along the root chord gives 74% on the LE and 26% on the TE.  Note that different airfoils and loading conditions will have different CP distributions, so don't use this approximation for any other purposes.  Also, this force split does not account for a tilting of the lift vector due to pulling angle of attack, but I will be considering this split sufficient for a static loading case.  This is certainly more representative than 50/50...

Note from the AVL plot that the root airfoils are working harder than the tip airfoils, with regard to a CP peak at the LE vs more evenly distributed along the chord.  Using the force distribution at the root (75/25) should provide a conservative estimate of the LE spar loading out toward the tips.  I may also run 100% on the LE just to see what that looks like.