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采用STAR-CCM+软件,用经过验证的湍流模型和方法,模拟空间(0~46km)环境中鼻锥模型在超声速飞行状态下的气动热特性,讨论了地面高焓风洞与空间飞行环境的区别.主要区别表现在如下3个方面:1要达到一定的滞止点温度,地面高焓风洞热环境依赖于超声速气动热与电弧加热的耦合作用;2在相同滞止温度的工况下,地面高焓风洞实验使得整个鼻锥实验件壁面都暴露在高温气流下,而空间飞行气动热主要集中在滞止点附近;3在相同滞止温度的工况下,空间飞行器的滞止区压力远远低于地面风洞实验压力.数值模拟揭示:相同来流马赫数下,随海拔高度增加,真实空间飞行条件下鼻锥滞止压力持续降低,而滞止温度则先下降然后升高;在同一空间高度下,随着来流马赫数增大,滞止温度和压力均呈抛物性增长,同时激波位置逐渐靠近前缘壁面,滞止区激波层变薄,但当来流马赫数高于4之后这种趋势将不再明显.
The aerodynamic thermal characteristics of the nose cone model under space flight (0 ~ 46km) under supersonic flight are simulated by STAR-CCM + software with a validated turbulence model and method, and the difference between the surface enthalpy wind tunnel and the space flight environment is discussed . The main differences are shown in the following three aspects: 1 To reach a certain stagnation point temperature, the thermal environment of the ground enthalpy wind tunnel depends on the coupling of supersonic aerodynamic heat and arc heating; 2 Under the same stagnation temperature conditions, The experiment of high enthalpy wind tunnel on the ground made the whole surface of the nose cone experiment exposed to high temperature airflow, while the aerodynamic heat of space flight mainly concentrated near the stagnation point. 3 Under the same stagnation temperature condition, The pressure is much lower than the experimental pressure of surface wind tunnel.Numerical simulation reveals that under the same Mach number, with the increase of altitude, the pressure of nose cone stagnation decreases continuously under real space flight conditions, while the stagnation temperature decreases first and then increases ; At the same space height, as the Mach number increases, the stagnation temperature and pressure all show a parabolic increase. At the same time, the shock position gradually approaches the front wall and the shock layer in the stagnant zone becomes thinner. However, Mach number is higher than After 4 this trend will no longer be obvious.