Abstract
Shock location inside the divergent section of convergent-divergent nozzles is studied experimentally. Pressure sensitive paint technique is used with six nozzles of different design Mach numbers in the range, 1.4-2.8. The results are corroborated by limited static pressure measurements. As it is well known, one-dimensional nozzle flow theory is found to grossly overpredict the throat-toshock-location distance at a given nozzle pressure ratio. A correlation from the literature based on rocket nozzle database is also found to be inadequate for prediction of the shock location especially for nozzles of lower design Mach number typical of aircraft application. For the parametric range covered with the set of nozzles used, a simple correlation is observed. All data collapse in a cluster when plotted as a function of the ratio of ‘jet Mach number’ to design Mach number; a curve-fit equation representing the average trend is provided.
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