Nozzle flow separation

Abstract

Flow separation in supersonic convergent–divergent nozzles has been the subject of several experimental and numerical studies in the past. Today, with the renewed interest in supersonic flights and space vehicles, the subject has become increasingly important, especially for aerospace applications (rockets, missiles, supersonic aircrafts, etc.). Flow separation in supersonic nozzles is a basic fluid-dynamics phenomenon that occurs at a certain pressure ratio of chamber to ambient pressure, resulting in shock formation and shock/turbulent-boundary layer interaction inside the nozzle. From purely gas-dynamics point of view, this problem involves basic structure of shock interactions with separation shock, which consists of incident shock, Mach reflections, reflected shock, triple point and sliplines (see Figs. 1, 2). Several viscous phenomena, such as boundary layers with adverse pressure gradients, induced separation, recirculation bubbles, shear layers may additionally occur and can strongly affect the flow-field inside the nozzle (see Figs. 3, 4). Previous studies on supersonic nozzles [1,2] have shown that shock-wave/boundary layer interaction (SWBLI) occurring in highly overexpanded nozzles may exhibit strong unsteadiness that cause symmetrical or unsymmetrical flow separation. In rocket design community, shock-induced separation is considered undesirable because an asymmetry in the flow can yield dangerous lateral forces, the so-called