Computation of internal high-speed separated flow with modified B–L and J–K models

The existence of shock–turbulent boundary layer interactions lead to very complicated flow phenomena and pose a challenge for numerical simulation. In this paper, two turbulence models, the Baldwin–Lomax (B–L) model and the Johnson–King (J–K) model, which were originally developed for simple external flow simulation, are modified to model complex high-speed internal separated flows. The full Navier–Stokes solver used in this paper is based on a cell-centered finite volume method and multistepping time marching scheme. Both implicit residual smoothing and local time stepping techniques are incorporated to accelerate the convergence rate. To ensure the numerical stability with the present explicit scheme, a point-implicit treatment to the source term in the ordinary differential equation (ODE) of the J–K model has been developed and has proved to be very effective in modeling such a complex flow. An arc-bump channel flow case has been studied. Comparisons of computed results with experimental data show that the present solver, with the modified turbulence models, predicts the shock and the flow separation very well. The J–K model is found to predict the size of the separation bubble with a higher accuracy. © 1998 John Wiley & Sons, Ltd.     

Nature of the surface heart transfer fluctuation in a hypersonic separated turbulent flow

This paper presents the results of an experimental study of the unsteady nature of a hypersonic separated turbulent flow. The nomimal test conditions were a freestream Mach number of 7.8 and a unit Reynolds number of 3.5×10⁷/m. The separated flow was generated using finite span forward facing steps. An array of flush mounted high spatial resolution and fast response platinum film resistance thermometers was used to make multi-channel measurements of the fluctuating surface heat trtansfer within the separated flow. Conditional sampling analysis of the signals shows that the root of separation shock wave consists of a series of compression wave extending over a streamwise length about one half of the incoming boundary layer thickness. The compression waves converge into a single leading shock beyond the boundary layer. The shock structure is unsteady and undergoes large-scale motion in the streamwise direction. The length scale of the motion is about 22 percent of the upstream influence length of the separation shock wave. There exists a wide band of frequency of oscillations of the shock system. Most of the frequencies are in the range of 1–3 kHz. The heat transfer fluctuates intermittently between the undisturbed level and the disturbed level within the range of motion of the separation shock wave. This intermittent phenomenon is considered as the consequence of the large-scale shock system oscillations. Downstream of the range of shock wave motion there is a separated region where the flow experiences continuous compression and no intermittency phenomenon is observed. The project supported by National Natural Science Foundation of China     

高超音速湍流分离表面热流率的脉动特性

本文给出高超音速湍流分离不稳定特性的实验研究结果。试验条件是:自由流马赫数为 7.8,单位长度雷诺数为 3.5×10~7/米。分离流场由有限展长前向台阶产生,并用有高空间分辨率和快速响应的一列平齐安装的铂膜电阻温度计和多通道系统测量其表面热流率脉动。信号的条件采样分析结果表明:分离激波的根部由一束压缩波构成,流向展长约二分之一来流边界层厚度,在边界层外汇聚成单一主激波。这种激波结构极其不稳定,出现大尺度运动,流向运动的尺度约为分离激波上游影响区域长度的22％。激波振荡频率为一宽频带,主要集中在 1～3 千赫。在分离激波运动区域,热流脉动呈间歇性,在无扰动和激波扰动间跳跃。可以认为这种间歇性是分离激波系统大尺度振荡的结果。在激波运动区域的下游为分离区,流体继续压缩,热流脉动无间歇。     

Numerical computations of supersonic inlet flow

Air‐breathing propulsion systems for high‐speed space travel applications are studied. Ramjets and scramjets have been identified as potential candidates. The flow inlets of such systems are modelled with a simulation that can predict all complex inlet flow features, including shock due to forebody, multiple shock reflections, normal shock, shock–boundary layer interaction and associated separation for two‐dimensional and axisymmetric inlets. Computed values are in good agreement with experimental data. Copyright © 2001 John Wiley & Sons, Ltd.     

Analysis of super compact finite difference method and application to simulation of vortex–shock interaction

Turbulence and aeroacoustic noise high‐order accurate schemes are required, and preferred, for solving complex flow fields with multi‐scale structures. In this paper a super compact finite difference method (SCFDM) is presented, the accuracy is analysed and the method is compared with a sixth‐order traditional and compact finite difference approximation. The comparison shows that the sixth‐order accurate super compact method has higher resolving efficiency. The sixth‐order super compact method, with a three‐stage Runge–Kutta method for approximation of the compressible Navier–Stokes equations, is used to solve the complex flow structures induced by vortex–shock interactions. The basic nature of the near‐field sound generated by interaction is studied. Copyright © 2001 John Wiley & Sons, Ltd.