Abstract
Shock-induced separation of turbulent boundary layers represents a long-studied problem in compressible flow, bearing, for example, on applications in high speed aerodynamics, rocketry, wind tunnel design, and turbomachinery. Experimental investigations have generally sought to expose essential physics using geometrically simple configurations.
Highlights
Shock-induced separation of turbulent boundary layers represents a long-studied problem in compressible flow, bearing, for example, on applications in high speed aerodynamics, rocketry, wind tunnel design, and turbo machinery
While a variety of computational and analytical methods have been developed for treating the problem, the methods are typically applicable to specific compressible flow regimes, i.e., transonic, supersonic or hypersonic flow, and due to the intrinsic unsteadiness of the separation process, require problem-specific tuning
This paper investigates time-average, shock-induced turbulent boundary layer separation in over-expanded rocket nozzles
Summary
Shock-induced separation of turbulent boundary layers represents a long-studied problem in compressible flow, bearing, for example, on applications in high speed aerodynamics, rocketry, wind tunnel design, and turbo machinery. The familiar rocket nozzle, known as a convergent-divergent, or de Laval nozzle, accomplishes this remarkable feat by simple geometry. In other words, it does this by varying the cross-sectional area (or diameter) in an exacting form. The analysis of a rocket nozzle involves the concept of “steady, one-dimensional compressible fluid flow of an ideal gas”. This means that: 1) The flow of the fluid (exhaust gases + condensed particles) is constant and does not change over time during the burn
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