A simplified method has been developed for design of supersonic wind tunnel nozzle which converts a subsonic flow into a uniform and parallel supersonic flow.
In order to obtain the initial value line from which calculation starts, transonic flow analysis is needed at the throat region. For the supersonic region, the hyperbolic type governing equations are solved by the conventional Method of Characteristics.
Since air is viscous, the development of boundary layer along the nozzle wall affects the actual nozzle contour and test-section Mach number is changed.
For the purpose of nozzle wall boundary layer correction the integral momentum and moment of momentum equation and skin friction law for the compressible turbulent boundary layer are adapted.
Boundary layer calculation begins at nozzle throat and initial conditions are given by Sibulkin & Merwin's empirical equations and the other boundary conditions (wall shape, boundary layer edge velocity, pressure, temperature,....) along the nozzle wall are given from the results of invicid flow calculation.
Boundary layer displacement thickness is used as a nozzle correction parameter and iterative method is used to obtain the nozzle contour of a given test-section size.
The results are compared with experimental value and found to be in good agreement.