In this paper , some flows around naca 0012 aerofoil , m6 wing and two different blunt vehicles are simulated , the results show that the present algorithm possesses good characteristics and capability for normal mach number and for hypersonic of simple configurations 通过对naca0012翼型、 m6机翼、钝头以及双钝锥?柱体的算例验证，本文能很好的解决一般马赫数问题，能够很好的求解简单外形的高超声速问题，对于复杂外形飞行器还需要进一步进行研究验证。
In this thesis , based on pershing ii surface to surface missile , a new kind of ballistic missile was designed with an additional rocket engine , which can be ignited twice . firstly , aerodynamic computational models of missile body and warhead which reentry with supersonic are built according to the task requirements ; secondly , the propulsion system model of missile is built whose first two stages are solid rocket engines and the third stage is liquid - solid combined rocket engine . the nozzle and the shape of the engine are designed to meet the needs of the populsion project ; thirdly , the trajectory model of the mass point is built and a wavy trajectory is designed & optimized ; finally , the ability of a missile ' s breaking through defence is analyzed 以美国潘兴导弹为原型，增加可两次点火的末级发动机，改装成具有跳跃能力的地地弹道导弹；首先，根据任务需求，建立了导弹的气动模型，并建立了弹头再入时高超声速气动模型；其次，建立了导弹推进系统模型，前两级采用了固体火箭发动机，第三级采用了固?液组合火箭发动机，并在总体方案要求下，对发动机喷管和外形进行了设计；第三部分，建立了导弹质点弹道模型，设计了一条跳跃式弹道，并对跳跃式弹道进行了优化设计；最后，对导弹进行了突防能力分析，从分析的结果可以看出，跳跃式弹道的突防能力比常规的抛物线弹道要强。
The tests were conducted in the hypersonic low density wind tunnel at nominal test conditions of mach 16 , stagnation temperature 923k , stagnation pressure 1 . 40mpa and 7 . 30mpa . heat - transfer data were obtained on a hemisphere model , a sharp cone and a big blunt cone respectively by means of infrared thermal mapping techniques , that of a 0 . 5mm thickness blunt cone by virtues of thermocouples . furth ermore , heat - transfer on all those models was calculated with the theoretical method 最后在名义m _ = 16 、 t _ 0 = 923k 、 p _ 0 = 1 . 40mpa及7 . 30mpa的高超声速低密度风洞中，利用红外热图技术获得了半球圆柱、尖锥、大钝头三个模型表面热流分布，利用薄壁法技术得到了一壁厚为0 . 5mm的钝锥模型表面的热流分布，并通过工程理论方法计算了模型表面的气动热，把理论计算结果与上述试验结果比较，几者符合得较好。
Although a dual - mode scramjet ' s configuration is simple and mainly consists of inlet , combustor and wake nozzle , its working process is complicated , especially in the combustor , involving a lot of subjects , including hypersonic aerodynamics , combustion chemistry , etc . the inner flow of a combustor is three - dimensional and complicated , including the interaction of shock wave , deflagration , vortex and boundary layer , and so on 它涉及到高超声速空气动力学、燃烧化学、扩散传质等多门学科；其内部的实际流动是复杂的三维流动过程，充满着激波、膨胀波、燃烧波、各种涡系、附面层及其相互之间的干扰，因此，燃烧室问题是整个发动机研究的关键所在。
To prove the accuracy of the mach number , and the parameter homogeneity of the design nozzle " s exit , cfd calculate has carried on the design results . under the condition of supersonic and hypersonic flow , and a certain range of temperature , and mach number , the conclusion of the influence of specific heat to nozzle design is drawn 为了验证所设计的喷管出口马赫数的大小和喷管出口流场的均匀性，采用nnd格式和b l湍流模型求解雷诺平均n - s方程，对设计结果进行了cfd验算，得出了在一定温度范围内，超音速、高超声速流动的条件下，不同马赫数范围内变比热容对喷管型面和喷管出口马赫数的影响。
Lateral jet control technology is researched in this dissertation , based on supersonic / hypersonic missile aerodynamics and lateral jet interaction ( ji ) effects associated with the attitude control solid - propellant rocket motors system in the low endoatmospheric range 横向喷流干扰效应研究在超声速高超声速导弹总体设计和精确制导技术研究领域一直占有重要地位。本论文针对大气层内超声速高超声速导弹采用姿态固体火箭发动机侧喷流控制技术的一些问题进行了研究。
In this study a primary method for designing a waverider configuration is developed based on the approximate solution of conical flow fields . the simplified cone - derived shock wave is discussed as the basic model for design . different shapes designed from different compression angles and basic geometric coefficients at the same mach number are put into analysis 本文基于高超声速条件下锥型流近似解提出了高超声速乘波构形的初步设计方案，对于不同马赫数、压缩角以及几何设计参数进行了外形设计并就参数对设计外形的影响进行了讨论，得到了设计参数影响外形的基本规律。
