The hypersonic flow at orbital speeds is a fundamental issue for the ground tests of aerospace crafts.The detonation-driven high-enthalpy expansion tube(JF16 expansion tube)was developed to investigate re-entry physics.A forward detonation cavity(FDC)driver was applied in the JF16 expansion tube to create stable driving flows.The sound speed ratio of the detonated to test gas was examined to minimize the magnitude of test flow perturbations.The acceleration section length,incident shock decay and diaphragms thickness were investigated in detail to obtain optimal operation parameters.Flow visualization was also carried out with schlieren system to demonstrate the test flow stability and the effective test duration.Experimental data showed that the test flow with a velocity of 8.3 km/s and a total enthalpy up to 40 MJ/kg can be generated successfully and the test duration lasts for more than 50μs.
JIANG Zong LinWU BoGAO Yun LiangZHAO WeiHU Zong Min
Supersonic model combustors using two-stage injections of supercritical kerosene were experimentally investigated in both Mach 2.5 and 3.0 model combustors with stagnation temperatures of approximately 1,750 K. Supercritical kerosene of approximately 760 K was prepared and injected in the overall equivalence ratio range of 0.5-1.46. Two pairs of integrated injector/flameholder cavity modules in tandem were used to facilitate fuel-air mixing and stable combustion. For single-stage fuel injection at an upstream location, it was found that the boundary layer separation could propagate into the isolator with increasing fuel equivalence ratio due to excessive local heat release, which in turns changed the entry airflow conditions. Moving the fuel injection to a further downstream location could alleviate the problem, while it would result in a decrease in combustion efficiency due to shorter fuel residence time. With two-stage fuel injections the overall combustor performance was shown to be improved and kerosene injections at fuel rich conditions could be reached without the upstream propagation of the boundary layer separation into the isolator. Furthermore, effects of the entry Mach number and pilot hydrogen on combustion performance were also studied.
The present paper employs the direct simulation Monte Carlo (DSMC) method to study the Rayleigh-Bénard flows, where the temperature ratio of the upper to lower plate is fixed to 0.1. For a Knudsen number (Kn) of 0.01, as the Rayleigh number (Ra) increases, the flow changes from the thermal conductive state to the convective state at about Ra=1700, and the calculated relation of heat flux through the lower plate versus Ra is in good agreement with classical experimental and theoretical results. For Kn=0.05, the thermal conductive state remains stable, and the increase of Ra cannot trigger thermal instability.
Aeroheating prediction is a challenging and critical problem for the design and optimization of hypersonic vehicles.One challenge is that the solution of the Navier-Stokes equations strongly depends on the computational mesh.In this letter,the effect of mesh resolution on heat flux prediction is studied.It is found that mesh-independent solutions can be obtained using fine mesh,whose accuracy is confirmed by results from kinetic particle simulation.It is analyzed that mesh-induced numerical error comes mainly from the flux calculation in the boundary layer whereas the temperature gradient on the surface can be evaluated using a wall function.Numerical schemes having strong capability of boundary layer capture are therefore recommended for hypersonic heating prediction.
Quanhua Sun,~(a) Huiyu Zhu,Gang Wang,and Jing Fan Key Laboratory of High Temperature Gas Dynamics,Institute of Mechanics,Chinese Academy of Sciences, Beijing 100190,China
