The Calculation of Drag for Airfoil Sections and Bodies of Revolution at Subcritical Speeds

A method is developed for calculating the drag, in a real compressible fluid and at subcritical Mach numbers, of airfoil sections at arbitrary lift coefficients and of bodies of revolution at zero angle of attack. To apply the method it is necessary to know the velocity distribution for airfoils and the velocity and thickness distributions for bodies of revolution, together with the Mach number of the free-stream transition point from laminar to turbulent flow, and the Reynolds number based on chord or axial length. The method consists of tracing the growth of momentum thickness along the surface, for both the laminar and turbulent boundary layers, by means of relations which involve elementary integrals and can be evaluated by simple numerical means. An outline of the computational procedures required for drag calculations is presented in the appendix to the report. The values of drag coefficient, computed by the method of the present report for a number of cases, are compared with the values obtained for the same configurations by other methods and the differences between the various results are found to lie within the limits of accuracy of current experimental techniques. The use of the present method is recommended by its simplicity and generality.