Abstract
To analyze the effectiveness of the thermal assessment test, the simulation method of the ground test in the arc-heated wind tunnel is studied. Based on the solution of the thermochemical nonequilibrium Navier-Stokes equations, the flowfield around the spherical cylinder is simulated in the flight and ground test conditions, and the difference in the high enthalpy flowfield between the flight and ground test conditions is investigated. The flight parameters and ground test conditions are selected according to the criterion that the total enthalpy and the stagnation point heat flux of the fully catalytic cold wall (calibrated heat flux) are similar. The flowfield for different temperature boundaries and different catalytic walls is solved under the same free stream conditions, and the stagnation point heat flux and oxygen atom mass fraction are compared and analyzed. It is found that the heat flux on the fully catalytic wall for the radiation balance temperature boundary in the ground test is lower than that in the corresponding flight condition, but the difference is not obvious on the noncatalytic wall. In addition, the oxygen atom mass fraction after the shock wave in the ground test is higher than that in the corresponding flight condition. To make the stagnation point heat flux and oxygen atom mass fraction after the shock wave similar to those of the flight, the simulation method of the arc-heated wind tunnel test needs to be adjusted.
Highlights
Hypersonic vehicles have received growing attention in recent years
The results show that the difference between test data and numerical simulation results for the fully catalytic wall is larger than that for the noncatalytic wall
Enthalpy is mainly related to flight velocity, and the calibrated heat flux increases with the decrease of flight altitude under certain enthalpy
Summary
Hypersonic vehicles have received growing attention in recent years. Due to their extremely high flying speed, the kinetic energy of the free stream is converted into internal energy by shock wave compression, forming hightemperature gas with thousands or even tens of thousands of Kelvins. Gas molecules dissociate or even ionize, resulting in a strong nonequilibrium effect. These physical and chemical phenomena of hypersonic flight form large aerodynamic thermal load, acting on the aircraft surface for a long time, posing a severe challenge to the thermal protection material and structure of the aircraft. The traditional conservative design method cannot meet the increasing performance requirements, so more accurate analysis and evaluation of the aerodynamic thermal environment for aircraft is urgently needed
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