【作者】 李斌；

【导师】 黄辉先；

【作者基本信息】 湘潭大学 ， 控制理论与控制工程， 2013， 硕士

【摘要】 目前高超声速飞行器的研究已经逐渐步入高潮。高超声速进气道/隔离段为超声速冲压发动机燃烧室提供稳定高质量且适用于燃烧的气流，对实现高超声速飞行起着关键作用，其性能直接影响燃烧室的燃烧。本文先对数值模拟方法进行了阐述，给出了本文所使用的数值计算方法和软件，本文选用FLUENT软件，采用有限体积法、隐式的二阶迎风格式、密度求解器（Density Based），采用AUSM+格式，层流粘性系数采用Sutherland公式，湍流模型采用Spalart-Allmaras（S-A）模型；本文还介绍了所采用的数值计算的流程与方法，前处理采用GAMBIT软件进行几何建模和网格划分，用FLUENT软件进行数值计算及后处理结果输出。本文对不同来流情况下的进气道/隔离段进行了二维流场数值模拟，结果首先与NASA的试验数据进行对比，说明了本文所使用的计算程序对超声速进气道/隔离段流场是适用的。数值计算得到了在非均匀气流下超声速进气道/隔离段的流场及其性能，分析了不同来流马赫数及非均匀来流条件下其对进气道/隔离段的影响及改进。数值仿真结果表明马赫数提高会降低冲压发动进气道的性能，激波反弹角增大，隔离段入口处附面层分离区增大，非均匀来流也使其总压恢复系数下降，出口畸变升高，对冲压发动机进气道和隔离段的性能造成很大影响。本文通过对隔离段入口处的上下壁面的调整，并数值计算出改进后进气道性能参数的规律变化，得到最佳的改进型模型Model B，计算结果表明改进后的进气道/隔离段（Model B）在非均匀来流下流场更加稳定，抗扰动能力更强。在气动性能方面，虽然改进后的进气道/隔离段出口总压畸变指数及出口马赫数变化不大，但总压恢复系数提高13%，这使得发动机推力上升，油耗下降。最后，本文计算并分析了进气道在不同飞行攻角条件下的流场情况，结果表明Model B具有更好的攻角适应性。更多还原

【Abstract】 The development of Hypersonic vehicle has been pushed to a new climax.Hypersonic inlet/isolator supplies high quality air flow for supersonic ramjetcombustion chamber and plays a key role in the achievement of the Hypersonic flight.Its performance directly affects the combustion in the combustion chamber.This article describes the numerical simulation method. The used numericalcalculation methods and software of the in this article are given. We choose theFLUENT software. And using the finite volume method, implicit second-orderupwind format and density based solver, using AUSM+format. Laminar flow viscouscoefficient adopt the Sutherland formula, turbulence model adopts theSpalart-Allmaras(S-A) model. This paper also describes the numerical calculationprocess. Pre-treatment uses the GAMBIT software for geometric modeling andmeshing. FLUENT is used to numerical calculation and after-treatment and the resultsoutput.In this paper, the flow field in the inlet and isolator under different airflowcondition are investigated. The flow field and performance index of the inlet andisolator of two levels of conical compression on non-uniform airflow are obtained.The influence and improved method of the inlet and isolator are analyzed. At first thecomputerized data are compared with Experiment data from NASA. The resultsshows that the calculation procedures used in this paper is applicable to supersonicinlet/isolator flow. Numerical simulation`s result shows that the total pressurerecovery coefficient is decreased on high Mach and non-uniform airflow. The resultsof numerical simulation shows that the performance of ramjet inlet is reduced byincreasing of Mach number. The shock rebound angle and the isolator entranceboundary layer separation are increased. Distortion index of the exit of the isolator isincreased. The total pressure recovery coefficient is decreased by non-uniform flow.The distortion of isolator exports was increased. It causes great impact to theperformance of scramjet inlet and isolation.In this paper, adjusted the isolated section of the up and down the wall at theentrance and through numerical simulation the law of the change of the improvementof inlet performance parameters are obtained. The improved model B is obtained. Theresults shows that the flow field of improved inlet/isolated is more stable onnon-uniform flow field and has a stronger anti-disturbance ability. Although the export total pressure distortion index and Mach number ofinlet/isolator have a little change. The total pressure recovery coefficient is increased13%.The performance of the inlet is improved. The thrust of the engine is increasedand the fuel consumption is decreased. At last this article calculates and analyzes theflow field on different attack angle. The result shows that Model B has a betteradaptability of attack angle.更多还原