I will keep saying it. Lift unavoidably MUST go to zero at the tips. Any lift distribution with lift at the tip is exaggerating the shear and bending moment, which you then design for. Any lift distribution with lift at the tip is also exaggerating the Cl curve as you go outward, which you then make (un-needed) changes to correct for it. Relying upon a bogus lift distribution to make fine detail changes is bogus too. Lift distribution based entirely upon planview area is putting you there. Please get beyond it. These math models are all for the dual purposes of estimating loads (for structural design) and estimating Cl distributions along the span (to anticipate and correct stall progression). They need to as accurately as possible reflect reality or you will just be putting in unneeded changes that otherwise hinder utility, speed and efficiency.

Lift distribution usually looks elliptical and is an excellent starting point. Going further, the pressure distribution along the span is what is usually really close to elliptical. If you want to be a philosopher about this, yes, the pressures are distributed over the area, so you can start with spanwise elliptical in one column of your spreadsheet, then in the next column put in the fraction of root chord you have at each station, and multiply the local elliptical by that fraction. This is modifying the lift outboard based upon plan view. This reduces total lift and lowers the curve of shear and moment vs spanwise position. Yes, you will have to adjust your calculated stall speed upwards to carry the weight of the airplane and the downforce from the tail. That will avoid finding out later that your actual stall speed is higher than you thought...

You can go further. If you washout the wing (for whatever reason), either by twist or foil choices at root and tip, you can figure out what the reduction in local Cl would be as you go outward, divide local Cl by base Cl, and multiply your local lift by that ratio. This too reduces the total lift and lowers the shear and moment curves.

Then you can do the downward deflected aileron and its adjusted Cl over the portion of the wing that is aileron influenced. The wing should remain intact when the aileron is deflected at Vne, so this is a load case you need to check for structure. How? Same way, figure out the ratio of Cldeflected divided by Clbase and multiply by that value. This increases shear and bending moment curves. You might be concerned about stall here too.

All of this drives one to say, but don't these reduce total lift? Hmm. Where your foil is clean (no control deflections) local Cl can not exceed stall Cl. And where your high lift devices are deflected, the local Cl can not exceed the stall Cl of that deflection. So when any part your local Cl exceeds Cl at stall, you need to scale back Cl everywhere. Reduce AOA everywhere. Easy way is to find the biggest delta between computed Cl (which we know is not actually available to the wing) and max CL, then divide that by your wing's lift slope (include correction for AR). That is the number of degrees you need to lower the AOA. Or if you are working only in Cl, just lower the entire Cl curve by that difference between computed and known max Cl...

I just added three columns to your spreadsheet to correct elliptical for taper, for twist, and then for aileron deflection. All make your math model more closely resemble what is actually happening, and none exaggerate the actual wing mechanical loads or tendency to tip stall.

Billski