Don't worry about the airfoil at the start. You need to focus first on getting the wing area right.
Do you have a speed requirement at altitude? For example, do you have to be able to maintain a spot over the ground in some expected headwind?
Your airplane will fly at CL=W/(q*S).
q=0.5rhoV^2 is determined by your flight condition. If you are free to choose V, then things are more complex.
You want to fly at CL for best L/D. For a given drag polar, that is a specific value of CL.
If you know your target CL, flight condition, and weight -- you can solve for the wing area S that will make it happen.
You want as much span as you can get -- do you have a span constraint? If not, you will be limited by structures. Keeping S constant, increasing span will shrink the chord and also the thickness of the airfoil. Such a thin spindly wing will not support the load.
Once you know chord and flight condition, you will know the cruise Reynolds number. You also know the cruise lift coefficient. For a straight, high aspect ratio wing, your sectional lift coefficient will be close to the cruise lift coefficient. You should start by selecting an airfoil with a design cl that matches your cruise CL.
This process is iterative -- you make a guess about the drag polar, calculate everything through, and then update your drag polar estimate based on the real wing area, the airfoil selected, etc. Then iterate again. A few times through and you'll converge on a good wing loading and airfoil for your point.