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If you only care about efficiency, you want the one that has the highest Cl/Cd at the design Cl, which may not be zero AoA. Determine desired Cl using required lift (weight plus tail download) in the following:

L = Cl * A * .5 * r * V^2

where L is lift, Cl is coefficient of lift, A is wing area, r is density and V is Velocity. Explanation is at:

https://www.grc.nasa.gov/www/k-12/airplane/lifteq.html

Look through your Cl/AoA charts at that Cl then get Cd using the associated AoA. Pick the airfoil with the highest value of Cl cubed / Cd squared for maximum endurance. Pick the highest CL/Cd using the associated AoA for maximum range.

If you only care about efficiency, you want the one that has the highest Cl3/Cd2 for endurance and highest Cl/Cd for range at the design Cl, which may not be zero AoA. Determine desired Cl using required lift (weight plus tail download) in the following:

L = Cl * A * .5 * r * V^2

where L is lift, Cl is coefficient of lift, A is wing area, r is density and V is Velocity. Explanation is at:

https://www.grc.nasa.gov/www/k-12/airplane/lifteq.html

Look through your Cl/AoA charts at that Cl then get Cd using the associated AoA. Pick the airfoil with the highest value of Cl cubed / Cd squared for maximum endurance. Pick the highest CL/Cd using the associated AoA for maximum range.

If you only care about efficiency, you want the one that has the highest Cl/Cd at the design Cl, which may not be zero AoA. Determine desired Cl using required lift (weight plus tail download) in the following:

L = Cl * A * .5 * r * V^2

where L is lift, Cl is coefficient of lift, A is wing area, r is density and V is Velocity. Explanation is at:

https://www.grc.nasa.gov/www/k-12/airplane/lifteq.html

Look through your Cl/AoA charts at that Cl then pick the highest CL/Cd using the associated AoA.

If you only care about efficiency, you want the one that has the highest Cl/Cd at the design Cl, which may not be zero AoA. Determine desired Cl using required lift (weight plus tail download) in the following:

L = Cl * A * .5 * r * V^2

where L is lift, Cl is coefficient of lift, A is wing area, r is density and V is Velocity. Explanation is at:

https://www.grc.nasa.gov/www/k-12/airplane/lifteq.html

Look through your Cl/AoA charts at that Cl then get Cd using the associated AoA. Pick the airfoil with the highest value of Cl cubed / Cd squared for maximum endurance. Pick the highest CL/Cd using the associated AoA for maximum range.

If you only care about efficiency, you want the one that has the highest Cl/Cd at the design Cl, which may not be zero AoA. Determine desired Cl using required lift (weight plus tail download) in the following:

L = Cl * A * .5 * r * V^2

where L is lift, Cl is coefficient of lift, A is wing area, r is density and V is Velocity. Explanation is at:

https://www.grc.nasa.gov/www/k-12/airplane/lifteq.html

Look through your charts at that Cl then pick the one with the highest CL/Cd. It will show you the associated AoA.

If you only care about efficiency, you want the one that has the highest Cl/Cd at the design Cl, which may not be zero AoA. Determine desired Cl using required lift (weight plus tail download) in the following:

L = Cl * A * .5 * r * V^2

where L is lift, Cl is coefficient of lift, A is wing area, r is density and V is Velocity. Explanation is at:

https://www.grc.nasa.gov/www/k-12/airplane/lifteq.html

Look through your Cl/AoA charts at that Cl then pick the highest CL/Cd using the associated AoA.