In divergent duct , adding heat can substitute the usage of geometry structure whatever mach number is . from force and thermodynamic cycle aspects , the effect of every components on engine performance was discussed 结果认为，对于后体为纯扩张构型的引射火箭，不论发动机通道内的马赫数是多少，都可以通过加热(改变加热位置)代替常规几何构型所起的作用。
In the results of eight 2d simulation cases with convective mach number ranged from 0 . 2 to 1 . 0 the mean velocities and the thickness growth rates are in good agreement with experimental results , but the turbulent fluctuation intensities are larger 对对流马赫数从0 . 2到1 . 0的混合层的二维数值研究表明，平均速度分布和扩张率与实验结果相符合，但湍流脉动强度偏大。
With present broadband spectrum forcing added on the inflow there still exist large vortex pairing processes in the mixing layer with high convective mach numbers , although the rolls up of large vortex need much longer spatial developing distance 在本文的入口宽带扰动下，高对流马赫数下二维混合层中仍然存在大涡的对并过程，只是大涡的卷起需要更长的空间发展距离。
The relevance functions between total pressure recover coefficient and the parameters ( mach number , angle of attack , flow coefficient , angle of sideslip ) for subsonic submerged inlet in turbojet and turbofan engine were investigated 摘要根据某埋入式进气道试验数据建立了计算埋入式亚声速进气道总压恢复系数的经验公式，表现为关于马赫数、流量系数、攻角、侧滑角的函数。
In order to research the aerodynamic characteristics of grid fins effected by width to length ratio , the grid fins of three hight to width ratios were adapted in wind tunnel experiments at subsonic , transonic and supersonic speeds 为研究它对栅格翼气动特性的影响规律，本研究采用风洞实验的方法对3种不同格宽翼弦比的栅格翼在亚、跨、超声速3个马赫数下进行了气动特性的研究。
Aiming at the high control performance of the mach number system in nf - 6 wind tunnel , we introduce the self - adapt control strategy based on bp neural network ( nn ) , which it use adapt character of nn to control mach number of nf - 6 wind tunnel 针对nf - 6风洞对马赫数控制系统提出的高要求，本文采用了一种基于bp神经网络的自适应控制方法，即：用神经网络自适应特性对风洞马赫数进行控制。
In the range of flight mach number mo = 3 ~ 7 , in order to take the advantage of ramjet and scramjet , a fix - shaped dual - mode scramjet , in which both subsonic and supersonic combustion can be obtained , is used for hypersonic vehicle 在飞行马赫数m _ = 3 - 7的范围内，为充分发挥冲压发动机的性能优势，高超声速飞行器采用了既能在亚声速燃烧模态，又能在超声速燃烧模态下工作的固定几何形状的双模态冲压发动机。
The full flow field characteristics of a supersonic inlet are illustrated at design conditions ; the influence of attack angle on flow field of the inlet is investigated ; b - spline theory is used to model the inlet with attack angle 给出了设计状态下进气道内外流场特征；分析了攻角变化对进气道流场的影响；以数值仿真结果为基础，利用b样条理论建立了反映攻角、马赫数及可调斜板角度变化的超声速进气道数学模型。
The results prove that : ( 1 ) when the mach number of the flow at the exit increases , the total pressure recovery decreases , and the circular steady total pressure distortion coefficient , turbulence intensity and synthesis distortion increase 研究结果表明： ( 1 )地面工作状态下，随着出口马赫数的增加，蛇形进气道出口截面的总压恢复系数不断下降，而稳态周向畸变指数、紊流度和综合畸变指数均上升，稳态径向畸变指数变化不大。
In order to accelerate the convergence speed , we employ the local time step , and implicit residual smoothing methods . finally , we use this method to get the steady solution of the flow around naca0012 and naca4412 airfoil at very low speed and some basic unsteady solution 本文对在低马赫数下绕naca0012翼型和naca4412翼型的定常流场进行了求解，结果和实验基本吻合，并对非定常运动情况进行了初步模拟研究，得出了一些有意义的结果。
