Hypersonic shock wave and turbulent boundary layer interaction in a sharp cone/flare model

A direct numerical simulation of hypersonic Shock wave and Turbulent Boundary Layer Interaction(STBLI)at Mach 6.0 on a sharp 7.half-angle circular cone/flare configuration at zero angle of attack is performed.The flare angle is 34.and the momentum thickness Reynolds number based on the incoming turbulent boundary layer on the sharp circular cone is Re θ=2506.It is found that the mean flow is separated and the separation bubble occurring near the corner exhibits unsteadiness.The Reynolds analogy factor changes dramatically across the interaction,and varies between 1.06 and 1.27 in the downstream region,while the QP85 scaling factor has a nearly constant value of 0.5 across the interaction.The evolution of the reattached boundary layer is characterized in terms of the mean profiles,the Reynolds stress components,the anisotropy tensor and the turbulence kinetic energy.It is argued that the recovery is incomplete and the near-wall asymptotic behavior does not occur for the hypersonic interaction.In addition,mean skin friction decomposition in an axisymmetric turbulent boundary layer is carried out for the first time.Downstream of the interaction,the contributions of transverse curvature and body divergence are negligible,whereas the positive contribution associated with the turbulence kinetic energy production and the negative spatial-growth contribution are dominant.Based on scale decomposition,the positive contribution is further divided into terms with different spanwise length scales.The negative contribution is analyzed by comparing the convective term,the streamwise-heterogeneity term and the pressure gradient term.

钝度对钝锥/楔熵层稳定性的影响研究

本文利用稳定性理论和数值模拟方法研究头部钝度对高超声速钝锥与钝楔的熵层稳定性的影响.结果表明钝楔的熵层不稳定性相比钝锥更强.随着钝度增加,熵层厚度逐渐变厚,熵层模态的不稳定性频率逐渐减小,但其不稳定区间不断向下游拓宽,表明熵层中的扰动将有更多机会影响边界层的转捩过程.数值模拟实现了熵层模态激发边界层第一模态的过程.

Predicting near-wall turbulence with minimal flow units in compressible turbulent channel flows

Mach number effects on the near-wall turbulence in the absence of outer motions remain unclear so far.The present study extends the Minimal Flow Units(MFUs),a widely applied method to investigate near-wall turbulence free from the impact of large-scale motions in the outer region in incompressible channel flows,to compressible wall-bounded turbulence.The compressible near-wall turbulence in MFU proves accurate in replicating near-wall statistics,independent of Mach number and statistically equivalent to the universal signals extracted from the full-sized channel.It is further utilized as universal signals in the predictive models of compressible near-wall turbulence,which is capable of accurately predicting variances and joint probability density functions of velocity and temperature fluctuations.

Progress in flight tests of hypersonic boundary layer transition

Boundary layer transition(BLT)can cause a sharp rise in heat flux and skin friction,which can seriously affect the flight performance and safety of hypersonic flight vehicles.Therefore,the mechanism,prediction and control of transition have become important issues that must be dealt with for the development of advanced flight vehicles,and it is also a research hotspot of particular interest to major aerospace countries.Compared to other transition research approaches,model flight tests can better present the transition problems under real flight conditions,thus have been carried out extensively over the past 30 years.The United States,Germany,France,Australia,and other countries have carried out transition research based on flight tests,such as the Pegasus wing-glove crossflow transition and the Hypersonic Boundary Layer Transition(HyBOLT)transition control flight test of the United States,the joint research project of the Hypersonic International Flight Research and Experimentation-1(HIFiRE-1)circular cone and the HIFiRE-5 elliptic cone transition flight tests between the United States and Australia,the flight test of compression surface transition of the scramjet forebody(LEA)in France and so on.Although these flight tests suffered various setbacks,they still obtained valuable transition data.Recently,the United States is carrying out the concave-surface transition flight tests of Hypersonic Boundary Layer Transition(BOLT)and BOLT-II.Since its first model flight test mission for verification purpose launched successfully in 2015,several hypersonic BLT flight tests have been conducted by China Aerodynamics Research and Development Center(CARDC).The flight tests have measured valid transition data under flight conditions,obtained the transition front and its dynamical variation on blunt cones at various angles of attack and a lifting body Hypersonic Transition Research Vehicle(HyTRV).The crossflow traveling waves in high-altitude flight were measured for the first time,and our understanding of hypersonic BLT has been greatly improved.

Design and transition characteristics of a standard model for hypersonic boundary layer transition research

To understand fundamental problems in hypersonic laminar-turbulent boundary layer transition for three-dimensional complex vehicles,a new standard model with typical lifting-body features has been proposed,named as hypersonic transition research vehicle(HyTRV).The configuration of HyTRV is fully analytical,and details of the design process are discussed in this study.The transition characteristics for HyTRV are investigated using three combined methods,i.e.,theoretical analyses,numerical simulations,and wind tunnel experiments.Results show that the fully analytic parameterization design of HyTRV can satisfy the model simplification requirements from both numerical simulations and wind tunnel experiments.Meanwhile,the flow field of HyTRV reveals typical transition mechanisms in six relatively separated regions,including the streamwise vortex instability,crossflow instability,secondary instability,and attachment-line instability.Therefore,the proposed HyTRV model is valuable for fundamental researches in hypersonic boundary layer transition.

Wall pressure beneath a transitional hypersonic boundary layer over an inclined straight circular cone

Properties of wall pressure beneath a transitional hypersonic boundary layer over a 7°half-angle blunt cone at angle of attack 6°are studied by Direct Numerical Simulation.The wall pressure has two distinct frequency peaks.The low-frequency peak with f≈10−50 kHz is very likely the unsteady crossflow mode based on its convection direction,i.e.along the axial direction and towards the windward symmetry ray.Highfrequency peaks are roughly proportional to the local boundary layer thickness.Along the trajectories of stationary crossflow vortices,the location of intense high-frequency wall pressure moves from the bottom of trough where the boundary layer is thin to the bottom of shoulder where the boundary layer is thick.By comparing the pressure field with that inside a high-speed transitional swept-wing boundary layer dominated by the z-type secondary crossflow mode,we found that the high-frequency signal originates from the Mack mode and evolves into the secondary crossflow instability.

From primary instabilities to secondary instabilities in Görtler vortex flows

We have studied the transformation process from primary instabilities to secondary instabilities with direct numerical simulations and stability theories(Spatial Biglobal and plane-marching parabolized stability equations)in detail.First Mack mode and second Mack mode are shown to be able to evolve into the sinuous mode and the varicose mode of secondary instability,respectively.Although the characteristics of second Mack mode eventually lose in the downstream due to the synchronization with the continuous spectrum,second Mack mode is found to be able to trigger a sequence of mode resonations which in turn give birth to highly unstable secondary instabilities.In contrast,first Mack mode does not involve in any mode synchronization.Nevertheless,it can still“jump”to a sinuous mode of secondary instability with a much larger growth rate than that of the first Mack mode.Therefore,secondary instabilities of Görtler vortices are highly receptive to the primary instabilities in the upstream,so that one should consider the primary instability in the upstream and the secondary instability in the downstream as a whole in order to get an accurate prediction of the boundary layer transition.

Experimental study of bluntness effects on hypersonic boundary-layer transition over a slender cone using surface mounted pressure sensors

In this work,we studied the bluntness effect on the hypersonic boundary-layer transition over a slender cone at Mach 6 with interchangeable tips in a noisy Ludwieg tube tunnel before the so-called“transition reversal”phenomenon occurs.The evolution of instability waves is characterized using surface flush-mounted pressure sensors deployed along the streamwise direction within unit Reynolds number from 4E+6/m≤Reunit≤10E+6/m,and the bluntness of the cone nose ranges from 0.1 mm to 5 mm.Power spectral density(PSD)of pressure fluctuation indicates that small nose bluntness(ReR≤2000)has little influence on the evolution of instability waves along the hypersonic boundary-layer,whereas with a moderate nose size(2000≤ReR≤5000),the hypersonic boundary layer transition is delayed monotonically as the nose radius increases before the boundary-layer turns into fully laminar without instability waves.The delaying effect can be attributed to the increased entropy-layer swallowing distance with a large tip radius.Instability wave characterization reveals that the second mode instability wave plays a dominant role before the transition reversal happens.The quadratic phase locking of second mode instabilities can be identified by bispectral analysis,and it attenuates as the nose tip radius increases.

可压缩壁湍流近壁解析大涡模拟的混合数值方法

本文提出了一种适用于可压缩壁湍流的近壁解析大涡模拟的混合数值方法.该方法基于壁湍流层级结构的特点,分别采用低耗散的动能守恒格式(KEP)和迎风格式求解近壁与外区的对流项,以保证准确分辨近壁的小尺度湍流脉动,并使外区的数值振荡引起的混淆误差对统计结果影响最小.该方法中存在两个待定参数,分别是数值格式的切换高度与切换范围参数.经过一系列数值实验,我们发现当切换高度与流向网格尺度一致时,统计结果与直接数值模拟的结果吻合最好.切换范围参数对结果影响不明显.本文进一步将所提出的方法应用在高雷诺数和高马赫数的壁湍流的数值模拟中,结果表明该方法在保持数值稳定的同时,可以准确计算壁面摩阻、平均速度和温度分布以及湍流脉动强度.

Effect of gas blowing on the head of thermal ablation vehicle

The reentry vehicle will encounter thermal ablation,especially at the stagnation point regime.A theoretical work has been done to analyze the thermal effect of gas blowing due to thermal ablation of surface material on the head of a general hypersonic vehicle.By deriving the formulation,research takes into account the effect of gas blowing on the thermal dynamics balance,and then solves them by numerical discretization.It is found that gas blowing will increase the temperature and heat flux at the surface of stagnation point regime.

A decomposition formula for the wall heat flux of a compressible boundary layer

Understanding the generation mechanism of the heat flux is essential for the design of hypersonic vehicles.We proposed a novel formula to decompose the heat flux coefficient into the contributions of different terms by integrating the conservative equation of the total energy.The reliability of the formula is well demonstrated by the direct numerical simulation results of a hypersonic transitional boundary layer.Through this formula,the exact process of the energy transport in the boundary layer can be explained and the dominant contributors to the heat flux can be explored,which are beneficial for the prediction of the heat and design of the thermal protection devices.