Chapter 4, Problem 4.15P

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 $9\times {10}^6$. Consider the Spitfire cruising at an altitude of 18,000 ft. (a) At what velocity is it flying for the root chord Reynolds number to be $9\times {10}^6$? (b) At this velocity and altitude, assuming completely turbulent flow, estimate the skin-friction drag coefficient for the NACA 2213 airfoil, and compare this with the total profile drag coefficient. Calculate the percentage of the profile drag coefficient that is due to pressure drag. Note: Assume that $\mu$ varies as the square root of temperature, as first discussed in Section 1.8.