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Basing on the solution of the Navier-Stokes equations and using the k-ω SST turbulence model, a numerical simulation of the supersonic flow in an air intake with an internal channel of a rectangular cross section was performed for the case when the shock wave generated by the inlet cowl interacted with the turbulent boundary layer on the inner surfaces of the channel. The three-dimensional configuration consisted of a convergent entrance section and a throat of a constant cross-sectional area is considered. Flow parameters at the entrance of the air intake were set taking into account the boundary layer developing on a pre-compression surface of an aircraft in front of the inlet. Two variants of the internal channel geometry were considered: firstly, when the lower surface of the channel was plane and, secondly, when this surface was arranged with longitudinal ribs. Numerical simulation of the three-dimensional flow was performed at the free-stream Mach number M = 4. [ABSTRACT FROM AUTHOR]

Numerical computations for studying the supersonic flows with a pseudo-shock wave were fulfilled. There was considered an axisymmetric duct including a funnel-shaped inlet, a throat section, a lengthened diffuser, and an outlet throttling section; the freestream flow with a Mach number M = 6 in front of the inlet was preset. There were applied the solvers including the Reynolds-averaged Navier-Stokes code and k-ω SST turbulence model for axisymmetric flows. The main subject of the investigation was determination of effects of the diffuser opening angle θd and the throttling level on forming the flow, gasdynamic and turbulent characteristics of the pseudo-shock wave. Data on variation of the turbulence viscosity along the duct length have shown the possibility of determination of the pseudo-shock wave termination in the diverging ducts instead of the usual method of determination of the pseudo-shock wave length by the distribution of the static pressure along the duct wall which is suitable just only for the non-diverging ducts. [ABSTRACT FROM AUTHOR]

Results of numerical computations of supersoni flow patterns in axisymmetric inlets with the internal compression of the airstream captured by them are presented. The simplest inlets in the form of conical funnels were under consideration. The characteristic feature of supersonic flows in these inlets is the formation of an initial longitudinally-curved shock wave with intensity increasing downstream and toward the flow axis, which is finalized by the generation of a triple-shock configuration including the initial incident shock wave, a central Mach disk, and a reflected shock wave. Conditions of forming these flows were investigated depending on the freestream Mach number, angle of the internal cone wall, and relative cross-sectional area of the inlet throat. A range of the freestream Mach number M = 2-12 was covered. There were determined the throat area required for the inlet starting and that minimal throat area at which the supersonic inflow into the started inlet still possible. An analysis of starting conditions of some axisymmetric inlets of internal compression tested at M = 7.5-11.5 in hot-shot wind tunnels was fulfilled. Firstly, data of these inlets tests were compared with data of numerical computations of flows in them carried out with the use of the pseudo-unsteady method for simulation of inviscid, laminar, and turbulent flows. Secondly, there was solved a problem on evolution of the unsteady turbulent flow forming in a common process of starting a hot-shot wind tunnel and a funnelshaped inlet tested in it. [ABSTRACT FROM AUTHOR]