Steady transonic flow through two-dimensional gas turbine cascades is efficiently predicted using a time-dependent formulation of the equations of motion. An integral representation of the equations has been used in which subsonic and supersonic regions of the flow field receive identical treatment. Mild shock structures are permitted to develop naturally without prior knowledge of their exact strength or position. Although the solutions yield a complete definition of the flow field, the primary aim is to produce airfoil surface pressure distributions for the design of aerodynamically efficient turbine blade contours. In order to demonstrate the accuracy of this method, computed airfoil pressure distributions have been compared to experimental results.

The flow field within the rotor of a transonic axial compressor has been computed and compared to measurements obtained with an advanced laser velocimeter. The compressor was designed for a total pressure ratio of 1.51 at a relative tip Mach number of 1.4. The comparisons are made at 100 percent design speed (20,260 RPM) with pressure ratios corresponding to peak efficiency, near surge, and wide open discharge operating conditions. The computational procedure iterates between a blade-to-blade calculation and an intrablade through flow calculation. Calculated Mach number contours, surface pressure distributions, and exit total pressure profiles are in agreement with the experimental data demonstrating the usefulness of quasi three-dimensional calculations in compressor design.

for a full three-dimensional boundary layer calculation with shock wave interaction. The development of such a complex computational model will require additional laser velocity measurements. These measurements should focus increased attention on the boundary layers and wakes. Data must be acquired on different fans at a variety of operating conditions.