Calculate the skin friction drag coefficient on the NACA 2415 airfoil, and compare your result with the experimental section drag coefficient in App.
In reality, the boundary layer on the airfoil discussed in Prob. 5.37 is neither fully laminar nor fully turbulent. The boundary layer starts out as laminar, and then transitions to turbulent at some point downstream of the leading edge (see the discussion in Sec. 4.19.) Assume that the critical Reynolds number for transition is 650,000. Calculate the skin friction drag coefficient on the NACA 2415 airfoil, and compare your result with the experimental section drag coefficient in App. D. Note: You will find from the answer to this problem that 86 percent of the airfoil section drag coefficient is due to skin friction and 14 percent due to pressure drag from :flow separation. Comparing this answer with the result of Prob. 5.36, which pertains to a thinner airfoil, we find that the pressure drag is a higher percentage for the thicker airfoil. However, for airfoils in general, the pressure drag is still a small percentage of the total drag. This drag breakdown is somewhat typical for airfoils at small angles of attack. By intent, the streamlined shape of airfoils results in small pressure drag, typically on the order of 15 percent of the total drag. Prob. 5.37 Here we continue in the vein of Probs. 5.34-5.36, except we examine a thicker airfoil and look at the relative percentages of skin friction and pressure drag for a thicker airfoil. Estimate the skin friction drag coefficient for the NACA 2415 airfoil in low-speed incompressible :flow at Re = 9 x 106 and zero angle of attack for (a) a laminar boundary layer, and (b) a turbulent boundary layer. Compare the results with the experimentally measured section drag coefficient given in App. D for the NACA 2415 airfoil. What does this tell you about the relative percentages of pressure drag and skin friction drag on the airfoil for each case? Probs. 5.34 Consider an NACA 2412 airfoil in a low-speed :flow at zero degrees angle of attack and a Reynolds numberof8.9 x 106. Calculate the percentage of drag from pressure drag due to :flow separation (form drag). Assume a fully turbulent boundary layer over the airfoil. Assume that the airfoil is thin enough that the skin-friction drag can be estimated by the :flat-plate results discussed in Ch. 4. Probs. -5.36 Returning to the conditions of Problem 5.34, where the boundary layer was assumed to be fully turbulent, let us now consider the real situation where the boundary layer starts out as laminar, and then makes a transition to turbulent somewhere downstream of the leading edge. Assume a transition Reynolds number of 500,000. For this case, calculate the percentage of drag that is due to :flow separation (form drag).
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