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跨声速计算
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  “跨声速计算”译为未确定词的双语例句
     The Calculation of Transonic Flow for Axis-Symmetric Euler Equations near M_∞=1
     M_∞=1附近轴对称Euler方程的跨声速计算
短句来源
     This formulation reserves the superiority of the classical potential formulation and extends its range of validity, and may be widely applied to the transonic flow calculations.
     这种方法既保持了势函数计算简便的优点,又扩大了势函数的适用范围,可以在跨声速计算中发挥很大的作用。
短句来源
     It is shown that the numerical integral steps that have got to convergence for chamber-nozzle flowfield computation are much more than those for transonic nozzle flowfield calculation.
     计算表明,达到收敛的数值积分步数比纯喷管的跨声速计算要多得多。
短句来源
  相似匹配句对
     We also give a new correction factor to simplify the calculation of the Unstatistic.
     的计算
短句来源
     Transonic Flow Simulation for Wing-Body Configurations
     机翼机身跨声速绕流的计算
短句来源
     The Parallel Computation of Transonic Flow
     分区并行跨声速流的计算
短句来源
     Calculation of Skewed Plates
     斜板的计算
短句来源
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Based on the theoretical analysis of the non-isentropic potential formulations which have been developed markedly in recent years, it is pointed out in the present paper that the formulation to solve the density directly from the momentum equations in the divergence from is more per-eecr. This formulation reserves the superiority of the classical potential formulation and extends its range of validity, and may be widely applied to the transonic flow calculations.

在对近年来有了很大发展的各种非等熵势函数方法进行理论分析的基础上,本文指出,从散度型动量方程直接求解密度并与势函数方程迭代的方法较为完善。这种方法既保持了势函数计算简便的优点,又扩大了势函数的适用范围,可以在跨声速计算中发挥很大的作用。

The chamber-nozzle subsonic-transonic flowfield of solid rocket motor is computed by time-dependent method. The governing equations are numerically solved by MacCormack explicit scheme. The parameters of boundary points are calculated with physical boundary conditions and characteristic equations on the reference plane. It is shown that the numerical integral steps that have got to convergence for chamber-nozzle flowfield computation are much more than those for transonic nozzle flowfield calculation. Although...

The chamber-nozzle subsonic-transonic flowfield of solid rocket motor is computed by time-dependent method. The governing equations are numerically solved by MacCormack explicit scheme. The parameters of boundary points are calculated with physical boundary conditions and characteristic equations on the reference plane. It is shown that the numerical integral steps that have got to convergence for chamber-nozzle flowfield computation are much more than those for transonic nozzle flowfield calculation. Although the Mach number distribution along the wall and axis for chamber-nozzle flowfield is similar to that for transonic nozzle flowfield, the iso-Mach number line distribution doesn't agree with the transonic nozzle flowfield.

本文用时间相关法计算了固体火箭发动机燃烧室─喷管亚跨声速流场数值解,控制方程用MacCormack二步显格式:边界点参数用物理边界条件和参考平面上的特征方程计算。计算表明,达到收敛的数值积分步数比纯喷管的跨声速计算要多得多。虽然喷管壁上和轴线上的马赫数分布与纯喷管计算类似,但喷管中的等马赫线分布与纯喷管计算的结果[3]相差较远。

Solutions of supersonic and transonic flows over sharp edged delta wings for moderate angles of attack are obtained numerically by solving the compressible Navier-Stokes equations using NND schemes with mixed flux splitting method.The results show that the vortex does not burst for the supersonic flow. The flow structure on the cross section perpendicular its axis depends on the sign of a parameter that is the derivative of ρuL. As the parameter is positive,the cross streamlines near the vortex core are inward....

Solutions of supersonic and transonic flows over sharp edged delta wings for moderate angles of attack are obtained numerically by solving the compressible Navier-Stokes equations using NND schemes with mixed flux splitting method.The results show that the vortex does not burst for the supersonic flow. The flow structure on the cross section perpendicular its axis depends on the sign of a parameter that is the derivative of ρuL. As the parameter is positive,the cross streamlines near the vortex core are inward. As it is negative, they are outward. If the parameter changes its sign along the axis, Hopf bifurcation takes place and results in a limit circle on the cross section. In the supersonic case, the parameter changes its sign once, a limit circle is found on the cross section. In the transonic case, the parameter changes its sign twice, there are two limit circles on the cross section.The vortex breakdown is obtained in transonic case. The vortex breakdown begins at spiral type and becomes bubble type in which there is a single vortex ring like structure.Numerical results show clearly the'dark hole' observed in experiment.

采用杂交通量分裂的NND格式模拟了跨、超声速三角翼背风区的旋涡运动结构。结果表明,旋涡的截面结构依赖于涡轴上的速度与密度的乘积沿轴向的导数λ=[(1/ρ)(ρu1/l)],当它大于零时,涡心附近的截面流线向里转,而当它小于零时,涡心附近的截面流线向外转,λ由大于零变到小于零,涡心附近将出现一个稳定的极限环,多一次变号,将多出现一个极限环。对超声速的计算,λ出现了一次变号,横截面流态中出现了一个极限环,对跨声速的计算,λ发生两次变号,横截面流态中出现了两个极限环。对跨声速三角翼的计算得到了涡破裂现象并发现涡破裂是从螺旋型起始并逐步演变成泡型。在破裂区,流动具有弱非定常特性。数值结果验证了张涵信的拓扑分析结果。

 
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