Abstract
The location of shock-induced boundary-layer separation inside the divergent section of convergent–divergent nozzles is studied experimentally. Pressure-sensitive paint technique is used with nozzles of different design Mach numbers in the range 1.4–2.8. Nozzle pressure ratios in the range of 1.12–4.91, corresponding to a “jet Mach number” range of 0.4–1.7 and a Reynolds number range of –, are covered in the experiment. As it is well-known, one-dimensional nozzle flow theory grossly overpredicts the throat-to-shock-location distance at a given nozzle pressure ratio. A correlation from the literature based on rocket nozzle databases is also found to be inadequate for these nozzles of lower design Mach number typical of aircraft applications. For the parametric range covered, a simple correlation for the shock location distance is found. 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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