For the conditions given in Problem 4.15, a more reasonable calculation of the skin friction coefficient would be to assume an initially laminar boundary layer starting at the leading edge, and then transitioning to a turbulent boundary layer at some point downstream. Calculate the skin-friction coefficient for the Spitfire’s airfoil described in Problem 4.15, but this time assuming a critical Reynolds number of 106 for transition.

Data from Problem 4.15: 

The airfoil section of the wing of the British Spitfire of World War II fame (see Figure 5.19) is an NACA 2213 at the wing root, tapering to an NACA 2205 at the wing tip. The root chord is 8.33 ft. The measured profile drag coefficient of the NACA 2213 airfoil is 0.006 at a Reynolds number of 9 × 106. Consider the Spitfire cruising at an altitude of 18,000 ft.