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Required Information

The alrfoll section of the wing of the British Spitfire of World War II fame is an NACA $2213$ at the wing root, tapering to an NACA $2205$ at the wing tip. The root chord is $8\mathrm{.}33$ ft $.$ The measured profile drag coefficient of the NACA $2213$ airfoll is $0\mathrm{.}006$ at a Reynolds number of $\backslash (9\backslash$times $10^\{6\}\backslash ).$ Consider the Spitfire crulsing at an altitude of $\backslash (17,000\backslash$mathrm$\{$ft$\}\backslash ).$ Assume that $\backslash (\backslash$mu $\backslash )$ varies as the square root of temperature.

At this velocity and altitude, assuming completely turbulent flow, estimate the skin$-$friction drag coefficient for the NACA $2213$ airfoll, and compare this with the total profile drag coefficient. Calculate the percentage of the profile drag coefficient that is due to pressure drag. $($Round the final answer to three decimal places.$)$

The skin$-$friction drag coefficient for the NACA $2213$ alrfoll is