The performance of supersonic turbine nozzles

Abstract

An investigation has been conducted at the National Gas Turbine Establishment into the performance of turbines having high pressure ratios per stage. The present Report discusses the mode of operation of supersonic nozzles for such turbines, and describes a cascade experiment. Both theory and experiment demonstrate that the conditions imposed upon the supersonic flow immediately downstream of the nozzles (e.g., by a following row of rotor blades) exert an overriding influence upon the nozzle outlet flow angle, and hence upon the maximum pressure ratio obtainable across the nozzle--providing that the axial component of velocity is subsonic. This is an important difference from the more familiar flow of subsonic turbine nozzles, where, for example, the downstream gas angle is controlled predominantly by the nozzle blade shape and spacing. A suitable test technique using a closed-jet tunnel is demonstrated. The particular nozzles tested, of convergent-divergent form, had a straight-sided divergent portion of 16½ deg total angle, a blade outlet angle of 76 deg (relative to axial direction) and a design Mach number of 2½. The flow was found to be well behaved as regards shock pattern, losses, and starting over the range of pressure ratios tested--between 9/1 and 19/1. In particular the efficiency at the design pressure ratio of 16-6/1 was high, the velocity coefficient calculated from traverses of pitot and static tubes being 0.98. For the conversion of pitot to total pressure at a Mach number of 2.5 a high accuracy is important in the measurement of the static pressure; nevertheless readings from a conventional four-hole instrument appear to be reliable.