Numerical Simulation of a Hypersonic Flow over an Aircraft in a High-Altitude Active Movement Area
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A hypersonic flow over an axisymmetric aircraft is numerically simulated in the case of a highly underexpanded exhaust plume (jet) of the main engine. The characteristics of the boundary layer separation occurring on the aircraft’s side surface are investigated for several successive points of its takeoff path. The Mach number at the nozzle exit is 6.5. The Mach number of the incoming flow varies from 4 to 7. In this case, the Reynolds number ranges from 2.5 × 105 to 3 × 103 and the ratio of the nozzle’s exit pressure to the ambient pressure, from 350 to 5 × 104. In the case of the Mach number of the incoming flow M∞ = 4, the variation range of the pressure ratio extends to 106. Replacement of the exhaust plume with a rigid simulator is considered. Data are obtained on the pressure ratios for which a separation flow begins to form on the side surface, the recirculation zone length, and the level of pressure in it in comparison with the available empirical dependences. A significant increase of the recirculation zone in front of the exhaust plume is shown when it is replaced by a rigid simulator of the same dimensions.
Keywordsnumerical simulation Navier-Stokes equations hypersonic flow underexpanded exhaust plume
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