【作者】 Syed Bilal Hussain Shah；

【导师】 陆夕云；

【作者基本信息】 中国科学技术大学 ， 流体力学， 2010， 博士

【摘要】 逆向流和横向流中的超声速喷射问题十分复杂,这类问题中存在射流与来流的相互作用及其复杂的激波振荡结构等。超声速射流在其逆向超声速来流或高超声速来流作用下,会形成弓形激波结构且具有振荡特性,因而被用于高超声速返回式飞行器的减阻。在超声速横向流作用下的声速射流可用于在低密度大气层中快速运行运载器的姿态控制,而在静止流中的射流还可用于推力矢量控制、火箭燃烧和流动混合的非设计工况下的不稳定侧向载荷的估算。本文数值模拟了在逆向流、横向流和静止流中的超声速、声速射流问题,文中着重研究了激波诱导的分离流动,探讨了射流干扰流场的机制。数值计算是采用PAK-3D的CFD程序实现,通过求解雷诺平均N-S方程(RANS),并采用合适的湍流模型,可以合理地模拟超声速和高超声速下的二维和三维的射流问题。通过对压比、流线、马赫数和入射角等参量的影响特性分析,给出了力和力矩综合效应的简化模型。同时,还详细分析探讨了逆向流、横向流和静止流中射流问题的激波结构和流动机理。首先,研究了半球柱体高超声速绕流中头部的逆向超声速射流问题。采用两方程湍流模型求解非定常雷诺平均方程,模拟了四种不同压比下高超声速来流问题。研究工作主要关注两种不同的典型流动形态,即SPM(短穿透模式)和LPM(长穿透模式),以及达到最大减阻效果的LPM。对于半球柱体结构外形,振荡流动形态SPM→LPM在所研究的压比流动中占主导。对常压比下的较高马赫数流动,获得了明显的减阻效果,可能是由于喷射的相对质量流量的增大所致。研究表明,倘若采用不同自由来流作用下的逆向喷射方式,其阻力降低规律可表示成质量流量的函数。进而,文中给出了不同自由来流下基于相对质量流量所描述的阻力减少量的简化数学模型。其次,研究了超声速横向流中的声速射流问题。通过一系列数值模拟,估算出随压比PR及攻角α增大时的喷射效率。运用半圆锥角10.4。钝锥体作为模型问题,在超声速流动中可以提供方向控制的效果。对马赫数4的自由来流,喷射与横向流的相互作用导致了净力的增加,通过改变喷射近前端和近尾端的压力分布,可以获得有利的俯仰力矩。进而,文中还提出了描述静态气动力系数随压比和攻角变化的数学模型。同时,基于激波结构、压力和马赫数的流场分布,对流动特性进行了分析。最后,研究了收缩-扩喷(CD)喷管的超声速射流问题。针对不同的面积收缩-扩张比,研究了“设计”和“非设计”两种工况下CD喷管的射流特性。对在设计工况下运行的CD喷管,发现研究结果与已公开发表的试验结果相符良好。此外,喷气舵置于CD喷管的超声速喷口,研究了喷气舵在0度偏转和20度偏转下的二种极端试验状态。对非设计工况,研究了不同的喷管压比NPR以及喷管面积比Ae／At,从1.2变化到1.6的流动特性。在较小的面积比和喷管压力比的条件下,喷管内的激波结构可视为正激波。在较小的面积比和中等的喷管压比条件下,可以观察到清晰的对称λ-激波结构。对于Ae／At=1.3和NPR≥1.70,可以观察到清晰的反对称λ-激波结构。反对称的λ-激波结构表明可运用射流的羽流来加强流动的混合效果。更多还原

【Abstract】 The supersonic jet exhausting in counterflow and crossflow is very complex problem owing to the nature of the jet interacting with the oncoming flow field and complex oscillatory shock structure. The supersonic jet exhausting in supersonic/hypersonic flows in counterflow direction has the oscillatory nature of the bow shock and is used for drag reduction of hypersonic vehicle. The sonic jet exhausting in the supersonic crossflow has the application in the attitude control for the fast moving vehicle in low density atmosphere whereas the jet exhausting in the quiescent flow has the application in thrust vector control, the estimation of unsteady side loads for off design conditions of rocket firing and flow mixing. The present study is aimed to apply the computational methodology for the shock induced separated turbulent flows and explore the mechanics in the jet interaction flow field. Specifically, the supersonic/sonic jet exhausting in the counterflow, crossflow and quiescent flow are investigated. The computational investigations are carried using a CFD code PAK-3D. The CFD code PAK-3D solve the Reynolds averaged Navier-Stokes (RANS) equations with a practical turbulence model to properly simulate the physics of supersonic and hypersonic two and three-dimensional jet interaction flowfield. The parametric study was performed to investigate the effect of pressure ratio, freestream Mach number and incidence. The mathematical models based on the calculations performed, were proposed defining the integrated effect in the form of force and moment. Further, the shock structures and flow physics for the three cases is discussed in detail.Firstly, a counterflow supersonic jet emanating from the nose of a hemispherical cylinder is investigated. Four different hypersonic oncoming freestream flows at various pressure ratios were solved using unsteady RANS equations with the two-equation k-εturbulence model. The investigations were mainly focused on two different flow regimes, SPM (short penetration mode)→LPM (long penetration mode) and LPM to ensure that the maximum drag reduction is achieved. For the present configuration being the hemispherical cylinder, the oscillatory flow regime SPM→LPM remains dominant over the most of the pressure ratios investigated. The obvious drag reduction is obtained for the higher Mach number flows at a constant pressure ratio, which can be attributed to the increase in relative mass flow rate of jet. The physical mechanism suggests that the drag reduction using a counterflow jet with different oncoming freestream flows can be expressed as a function of mass flow ratio. A model for the various freestream flows describing the drag reduction as a function of relative mass flow rate is proposed.Secondly, a sonic jet exhausting in the supersonic crossflow is investigated and series of numerical simulations were performed to evaluate the effectiveness of the jet with increase in pressure ratio and the angle of attack. The jet emanating from a blunted cone with a half-cone angle of 10.4°was used that provide possible directional control in supersonic flows. For an oncoming freestream flow of Mach number 4, the jet interaction in the crossflow causes an increase in net force and consequently a useful pitching moment is obtained through altering pressure distribution in its near forward and aft vicinity of the jet. A model describing the computed static aerodynamic coefficients variation with pressure ratios and angle of attack are presented to demonstrate this jet interaction effect. The shock structures, pressure and Mach number contours with streamline patterns gave a qualitative insight. For jet exhausting in the quiescent flow, the supersonic flow through various area ratio convergent-divergent, CD nozzles was investigated for both "designed "and "off designed" conditions. For CD nozzle operating at designed conditions, a reasonable agreement was found with the published experimental results. Further a jet vane was placed at supersonic exhaust of the CD nozzle. Two extreme test cases were studied with zero and 20°of jet vane deflection. For the off designed conditions, the nozzle area ratio Ae/At ranges from 1.2 to 1.6 with various nozzle pressure ratios NPR. At lower end of Ae/At and NPR, the shock structure inside the nozzle may be regarded as a normal shock. At lower Ae/At and moderate NPR, a well defined symmetricλ-shock was observed. For Ae/At = 1.3 with NPR≥1.70, a well defined asymmetricλ-shock structure was observed. The asymmetricλ-shock structure suggests that jet plume can be used for the mixing enhancement.更多还原