A method is developed for calculating the profile drag of a yawed wing of infinite span, based on the assumption that the form of the spanwise distribution of velocity in the boundary layer, whether laminar or turbulent, is insensitive to the chordwise pressure distribution. The form is assumed to be the same as that accepted for the boundary layer on an unyawed plate with zero external pressure gradient. Experimental evidence indicated that these assumptions are reasonable in this context. The method is applied to a flat plate and the NACA 64-012 section at zero incidence for a range of Reynolds numbers between 106 and 108. Angles of yaw up to 45°, and a range of transition point positions. It is shown that the drag coefficient of a flat plate varies with yaw as Cos½ ˄ (where ˄ is the angle of yaw) if the boundary layer is completely laminar, and it varies as Cos4/5 ˄ if the boundary layer is completely turbulent. The drag coefficient of the NACA 64-012 section, however, varies very closely as Cos˄ foe transition point positions between 0 and 0.5 c. Further calculations on wing sections of other shapes and thicknesses and more detailed experimental checks of the basic assumptions at higher Reynolds number are desirable.