Wing area (exposed, reference, gross) vs CL & Cl

[Mach Stall]

Here's a question for the highly technical aerodynamicists. Maybe I'll stump the band with this one.

My question is- between the various wing area definitions, which more accurately reflects the lift coefficients actually seen on the wing sections? Given that the difference in wing area definitions as relates, e.g., to a Boeing 787 is a whopping 25% in area (between Piano gross and exposed area), the question of definitions is hardly a minor distinction.

First, I think I understand the difference between exposed and reference wing area (reference area extends the imaginary wing planform trapezoid through the fuselage and does not include the "yehudi" / extensions; exposed is just like it sounds). Then there are gross area and "wimpress" definitions, which include the exposed yehudi or wing extensions and all or almost all of their imaginary extended portions which pass thru the fuselage.

And I think I understand the difference between CL and Cl (CL is for the whole wing, Cl is for a local airfoil).

One might intuitively think the exposed area is the most accurate since it is all that is hanging out there, but actually the "carry-through" wing area reflected in the other definitions' area is indirectly producing massive lift (via the fuselage) -- this is because the wing lift cannot simply stop on a dime at the junction of the fuselage (unless the fuselage were an infinite wall).

Another way to put my question is this: If someone at Boeing (pre-CFD days) were picking an airfoil for the 737, and the calculations showed that the lift coefficient at mid-span needed to be Cl 0.50 based on exposed area, would this more accurately select an airfoil than if they based the Cl on gross area (perhaps about Cl 0.40)?

Yet another way to put the question is -- were airfoils more accurately "labeled" by reference area, gross area, or exposed area re their design lift coefficient Cl? Let's say a designer (again, pre- CFD days) picks a laminar flow airfoil with a design lift coefficient of Cl 0.6 and puts it on a plane based on his gross area calculation. Will this airfoil be close to ideal, or should he have instead picked a Cl 0.75 airfoil because in flight exposed area is more accurate re Cl? And in the old days, not that long ago really, airfoils really were labelled with lift coefficients (eg, many NACA laminar flow series airfoils).

I realize the Cl varies considerably across the wingspan, depending on twist and 3D effects, but for any arbitrary point, my question still seems meaningful, albeit a tad nerdy.

Any thoughts (esp if someone were a designer)?