In this paper we investigate a technique of computing airfoil sections that permit shock-free transonic flow around them at a specific Mach number and angle of attack. By transonic we mean that the speed of the aircraft is less than the speed of sound, but close enough to it so that on top of the wing, where the airflow is fastest, the Mach number becomes greater than one. For air the Reynolds number is extremely high, and hence the viscous effects will be confined chiefly to a thin boundary layer provided separation effects can be suppressed. Therefore we take for our equations of motion the equations of potential flow.
KeywordsMach Number Singular Solution Complex Domain Physical Plane Transonic Flow
Unable to display preview. Download preview PDF.