Thermochemical non-equilibrium flow characteristics of high Mach number inlet in a wide operation range

The high-temperature non-equilibrium effect is a novel and significant issue in the flows over a high Mach number(above Mach 8)air-breathing vehicle.Thus,this study attempts to investigate the high-temperature non-equilibrium flows of a curved compression two-dimensional scramjet inlet at Mach 8 to 12 utilizing the two-dimensional non-equilibrium RANS calculations.Notably,the thermochemical non-equilibrium gas model can predict the actual high-temperature flows,and the numerical results of the other four thermochemical gas models are only used for comparative analysis.Firstly,the thermochemical non-equilibrium flow fields and work performance of the inlet at Mach 8 to 12 are analyzed.Then,the influences of high-temperature non-equilibrium effects on the starting characteristics of the inlet are investigated.The results reveal that a large separation bubble caused by the cowl shock/lower wall boundary layer interaction appears upstream of the shoulder,at Mach 8.The separation zone size is smaller,and its location is closer to the downstream area while the thermal process changes from frozen to non-equilibrium and then to equilibrium.With the increase of inflow Mach number,the thermochemical non-equilibrium effects in the whole inlet flow field gradually strengthen,so their influences on the overall work performance of the high Mach number inlet are more obvious.The vibrational relaxation or thermal non-equilibrium effects can yield more visible influences on the inlet performance than the chemical non-equilibrium reactions.The inlet in the thermochemical non-equilibrium flow can restart more easily than that in the thermochemical frozen flow.This work should provide a basis for the design and starting ability prediction of the high Mach number inlet in the wide operation range.

Design method for hypersonic bump inlet based on transverse pressure gradient

Transverse pressure gradient(TPG)is one of the key factors influencing the boundary layer airflow diversion in a bump inlet.This paper proposes a novel TPG-based hypersonic bump inlet design method.This method consists of two steps.First,a parametric optimization approach is employed to design a series of 2D inlets with various compression efficiencies.Then,according to the prescribed TPG,the optimized inlets are placed in different osculating planes to generate a 3D bump inlet.This method provides a means to directly control the aerodynamic parameters of the bump rather than the geometric parameters.By performing this method to a hypersonic chin inlet,a long and wide bump surface is formed in the compression wall,which leads to good integration of the bump/inlet.Results show that a part of the near-wall boundary layer flow is diverted by the bump,resulting in a slight decrease in the mass flow but a significant improvement in the total pressure recovery.In addition,the starting ability is significantly improved by adding the bump surface.Analysis reveals that the bump has a 3D rebuilding effect on the large-scale separation bubble of the unstarted inlet.Finally,a mass flow correction is performed on the designed bump inlet to increase the mass flow to full airflow capture.The results show that the mass flow rate of the corrected bump inlet reaches up to 0.9993,demonstrating that the correction method is effective.