Abstract
In rocket engine nozzle design, flow separation is essential and, given the high temperatures and pressures in the thrust chamber, regenerative cooling is critical for maintaining the nozzle wall's integrity. This passage provides a summary of an in-depth numerical analysis of boundary layer separation and heat transfer within a 30°-15° cooled nozzle. The performance of the SST-V turbulence model under these conditions is assessed numerically. A variety of factors are investigated, including wall temperature, turbulent Prandtl number, and constant specific heat ratios (spanning 1.31 to 1.4 for constant fluid properties of N2O, CH4, Cl2, and air). Furthermore, variable specific heat ratios (from 1.39 to 1.66 for variable fluid properties of air, CH4, O2, and Helium) are examined, along with other parameters that affect the location of separated flow and the local wall heat transfer.
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