Effect of disturbances at inlet on hypersonic boundary layer transition on a blunt cone at small angle of attack

To investigate the effect of different disturbances in the upstream, we present numerical simulation of transition for a hypersonic boundary layer on a 5-degree half-angle blunt cone in a freestream with Mach number 6 at 1-degree angle of attack. Evolution of small disturbances is simulated to compare with the linear stability theory （LST）, indicating that LST can provide a good prediction on the growth rate of the disturbance. The effect of different disturbances on transition is investigated. Transition onset distributions along the azimuthal direction are obtained with two groups of disturbances of different frequencies. It shows that transition onset is relevant to frequencies and amplitudes of the disturbances at the inlet, and is decided by the amplitudes of most unstable waves at the inlet. According to the characteristics of environmental disturbances in most wind tunnels, we explain why transition occurs leeside-forward and windside-aft over a circular cone at an angle of attack. Moreover, the indentation phenomenon in the transition curve on the leeward is also revealed.

Response of a hypersonic blunt cone boundary layer to slow acoustic waves with assessment of various routes of receptivity

The hypersonic boundary-layer receptivity to slow acoustic waves is investigated for the Mach 6 flow over a 5-degree half-angle blunt cone with the nose radius of 5.08 mm. The plane acoustic wave interacts with the bow shock, and generates all types of disturbances behind the shock, which may take various routes to generate the boundarylayer unstable mode. In this paper, two routes of receptivity are investigated in detail.One is through the disturbance in the entropy layer. The other is through the slow acoustic wave transmitted downstream the bow shock, which can excite the boundary-layer mode due to the synchronization mechanism. The results show that, for a low frequency slow acoustic wave, the latter route plays a leading role. The entropy-layer instability wave is able to excite the first mode near the neutral point, but its receptivity efficiency is much lower.

Accurately predicting hypersonic transitional flow on cone via a symmetry approach

A new algebraic transition model is proposed based on a Structural Ensemble Dynamics(SED)theory of wall turbulence,for accurately predicting the hypersonic flow heat transfer on cone.The model defines the eddy viscosity in terms of a two-dimensional multi-regime distribution of a Stress Length(SL)function,and hence is named as SED-SL.This paper presents clear evidence of precise predictions of transition onset location and peak heat flux of a wide range of hypersonic Transitional Boundary Layers(TrBL)around straight cone at zero incidence,to an unprecedented accuracy as validated by over 70 measurements for varying five crucial influential factors(Mach number,temperature ratio,cone half angle,nose Reynolds number and noise level).The results demonstrate the universality of the postulated multi-regime similarity structure,in characterizing not only the spatial non-uniform distribution of the eddy viscosity in hypersonic TrBL on cone,but also the dependence of the transition onset location on the five influential factors.The latter yields a novel correlation formula for transition center Reynolds number which takes similar functional form as the SL function within the symmetry approach.It is concluded that the SED-SL model simulates TrBL around cone with uniformly high accuracy,and then points out to an optimistic alternative way to construct hypersonic transition model.

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.

A symmetry-based length model for characterizing the hypersonic boundary layer transition on a slender cone at moderate incidence

The hypersonic boundary layer(HBL)transition on a slender cone at moderate incidence is studied via a symmetry-based length model:the SED-SL model.The SED-SL specifies an analytic stress length function(which defines the eddy viscosity)describing a physically sound two-dimensional multi-regime structure of transitional boundary layer.Previous studies showed accurate predictions,especially on the drag coefficient,by the SED-SL for airfoil flows at different subsonic Mach numbers,Reynolds numbers and angles of attack.Here,the SED-SL is extended to compute the hypersonic heat transfer on a 7∘half-angle straight cone at Mach numbers 6 and 7 and angles of attack from 0∘to 6∘.It is shown that a proper setting of the multi-regime structure with three parameters(i.e.a transition center,an after-transition near-wall eddy length,and a transition width quantifying transition overshoot)yields an accurate description of the surface heat fluxes measured in wind tunnels.Uniformly good agreements between simulations and measurements are obtained from windward to leeward side of the cone,implying the validity of the multi-regime description of the transition independent of instability mechanisms.It is concluded that a unified description for the HBL transition of cone is found,and might offer a basis for developing a new transition model that is simultaneously of computational simplicity,sound physics and greater accuracy.

Application of surrogate models to stability analysis and transition prediction in hypersonic flows

To increase the efficiency and robustness of stability-based transition prediction in flow simulations, simplified methods are introduced to substitute direct stability analyses for rapid disturbance growth prediction. For low-speed boundary layers, these methods are mainly established based on self-similar assumptions, which are not applicable to non-similar boundary layers in hypersonic flows. The objective of this article is to investigate the application of surrogate models to stability analysis of non-similar flows over blunt cones, focused on parameterization of boundary-layer (BL) profiles. Firstly, correlations between BL edge and profile parameters are analyzed, along with self-similar flow parameters and discrete points on BL profiles, which present four groups of BL characteristic parameters. Secondly, using these parameters as inputs, surrogate models are built for disturbance growth prediction over an MF-1 blunt cone. Results show that, surrogate models using four BL edge parameters and a BL shape factor {Ue, Te, ρe, ηe, H12} for stability analysis can achieve comparable accuracy with those using 16 discrete BL profile parameters, which are more precise than those using merely self-similar parameters or BL edge parameters. Thirdly, the established surrogate models are validated by stability analysis and transition prediction over the MF-1 blunt cone in flight experiments at the instants of t = 17 s ~ 22 s. Compared with direct linear stability analyses, the mean relative error of predicted disturbance growth rates by surrogate models is 8.0% and the maximum relative error of N factor envelopes is 6.6%, which indicates feasible applications of surrogate models to stability analysis and transition prediction of non-similar boundary layers in hypersonic flows.

7°尖锥高超声速边界层脉动压力实验研究

针对半锥角7°尖锥模型,在常规高超声速风洞中开展了边界层脉动压力测量实验,进行了线性稳定性分析,研究了单位雷诺数和马赫数对尖锥边界层转捩位置和边界层稳定性的影响规律。模型长度800 mm,头部半径0.05 mm,实验单位雷诺数0.49×10^(7)~2.45×10^(7) m^(-1),马赫数5~8,迎角0°。通过红外热图技术和高频脉动压力测量技术获得了模型表面边界层转捩位置和边界层内扰动波能谱分布,利用线性稳定性理论分析了最不稳定波频率和增长率。实验结果表明:在转捩区间可以测量到明显具有不稳定波频谱特征的脉动压力信号,其频率与稳定性理论分析的二模态不稳定波接近,幅值变化趋势也与之类似;随着雷诺数增大,不稳定波出现位置提前,频率增大,转捩位置提前;边界层中不稳定波包含第一和第二模态,马赫数5时,转捩由第一模态主导,马赫数高于6时,由第二模态主导。

钝度对高超声速钝锥边界层稳定性的影响

采用直接数值模拟方法计算了8个不同球头半径的钝锥基本流,运用线性稳定性理论分析了钝度对边界层稳定性的影响。结果表明,随钝度增大,边界层内的不稳定区向下游移动,第二模态的最大增长率减小。在线性稳定性分析的基础上,研究了非线性扰动演化以及平均流修正对稳定性的影响。结果表明,在基本流中引入有限幅值扰动后,下游的平均流剖面会发生明显改变。流场稳定性发生显著变化,线性阶段最不稳定的第二模态波变得稳定,而第一模态波明显增长起来。第一模态波的快速增长使N值可以达到4,这将会对转捩有很大的促进作用。

高超声速飞行器气动热辐射特性

飞行器在大气层内高速飞行时,高速气体来流在飞行器顶端形成高温高压气体绕流,并对顶端光学探测窗口形成强烈气动加热,光学窗口温度急剧上升,高温气体和光学窗口的红外辐射对探测系统形成严重的气动热辐射效应,探测信噪比下降。针对非灰混合气体和光学窗口材料的辐射特点,采用介质辐射传输方法,模拟了钝锥球头外形头部且顶端安装探测窗口飞行器的气动热辐射,研究了气体和窗口热辐射随时间的发展及其与窗口材料的关系。结果表明,高温气体的热辐射与飞行器的运动状态关系密切,而光学窗口的热辐射随飞行时间增加而迅速增强,逐渐成为气动热辐射的主要因素,因此抑制光学窗口的温度上升速度和幅度是减弱气动热辐射的关键。

Menter转捩模型在高超声速流动模拟中的改进及验证

通过压力梯度参数和湍流普朗特数的修正,对Menter等构造的低速经验关系式转捩模型进行了改进,将基于局部信息的Menter转捩模型用于高超声速流动转捩数值模拟。在多个高超声速尖锥转捩流动模拟中,对改进后模型的可靠性进行了检验,算例考核了不同壁面温度、来流湍流度、流向压力梯度等多种流动条件。结果表明,数值计算和风洞试验的壁面温度、温度恢复因子和热流等符合得较好,修正后模型较好地模拟了转捩起始位置和转捩区长度,在高超声速边界层转捩预测中具有一定潜力。

高超声速锥柱裙模型边界层转捩的弹道靶实验

为研究高超声速边界层转捩现象、给边界层计算提供可靠的对比数据，在中国空气动力研究与发展中心趟高速弹道靶上开展了锥柱裙模型高超声速边界层转捩的自由飞实验。所采用的锥柱裙模型全长105mm，飞行速度1．94km／s（Ma=5．65），单位雷诺数4．32×10^7～1．20×10^8m-1。使用激光阴影成像技术，获得了锥柱裙模型边界层转捩和湍流边界层发展的图像，测得的湍流边界层厚度在0．6～2．2mm之间，湍流涡的流向尺寸与边界层厚度的比值介于0．3～0．8之间且沿流向呈下降趋势。实验结果表明：弹道靶实验能够获得给定飞行环境下的高超声速边界层转捩图像，从图像中可以清晰判断转捩位置或区域、测量边界层厚度和分析湍流涡的尺寸。

小攻角高超声速钝锥边界层中不同扰动对转捩的影响

为了研究上游不同扰动对转捩位置的影响,针对来流Ma=6,攻角1°,半锥角5°的钝锥边界层的转捩进行了数值模拟.首先研究了边界层中小扰动的演化,与流动稳定性理论进行了对比,结果表明:在所考虑的流场中,流动稳定性线性理论可以对扰动的增长率做出一个较好的预测.在此基础上,研究了不同扰动对转捩位置的影响.计算给出了在两种不同频率分布的扰动情况下,转捩位置沿周向的分布.结果表明,转捩位置沿周向分布与入口扰动的幅值和频率有关.某子午面的转捩位置由该处的最不稳定波在入口的幅值决定.根据大多数风洞中背景扰动的特性,解释了小攻角圆锥转捩实验中背风面先转捩,迎风面后转捩的现象.同时,还解释了在背风面附近转捩位置"凹陷"的现象.

7°尖锥高超声速边界层转捩红外测量实验

为了推动高超声速边界层转捩研究的深入开展,给边界层转捩机理研究、物理模型验证、转捩数据库构建和转捩天地相关性的建立等提供基础风洞实验数据,在中国空气动力研究与发展中心的Φ1m高超声速风洞开展了边界层转捩规律红外热图实验。针对半锥角7°尖锥模型,研究了不同单位雷诺数、迎角和马赫数对尖锥边界层转捩位置的影响规律。实验单位雷诺数(0.49~2.45)×10~7/m,迎角范围-10°~10°,马赫数5~7,模型头部半径0.05mm。通过红外热图技术测量模型表面温度分布,获得了较为详细的转捩位置和转捩参数影响规律。实验结果表明:在马赫数5~7范围内,马赫数增大,尖锥转捩位置提前,分析认为是高马赫数时的雷诺数较大、自由流噪声水平较高引起;随着单位雷诺数的增大,边界层转捩位置前移,转捩雷诺数保持不变,约为3.0×10~6;小迎角时,随着迎角的增大,迎风面边界层转捩推迟,背风面边界层转捩前移,在10°大迎角时,迎风区中心线转捩前移,出现迎角"转捩逆转"现象,背风区出现了流动分离导致的低热流条带。

高超声速进气道前缘钝度效应试验研究

基于一种典型高超声速二元进气道,考察前缘钝度效应对进气道边界层转捩的影响,加工了四种半径为R=0.05mm,R=0.1mm,R=0.2mm,R=0.25mm的前缘,在FD-07风洞中开展了自然转捩及人工转捩的风洞试验。试验中采用压缩拐角压力分布特征及进气道起动相结合的方法来估计边界层转捩位置,得出了进气道压缩面边界层转捩位置随前缘半径变化的规律。试验表明在来流条件下随前缘钝化半径增加,边界层转捩位置明显后移。针对R=0.25mm时进气道不起动的情况,基于线性稳定性理论(LST)理论设计了人工转捩条带,通过试验成功实现了转捩。

钝锥绕流流动稳定性分析与转捩预报

研究了超音速钝锥绕流的稳定性和转捩点预报的数值计算方法 ,首先采用Euler方程求解钝锥绕流基本流场 ,用所得到的物面压力分布作为粘性边界层的外缘压力分布 ,给出了基本流场的初值 ;然后应用反迭代法与边界层渐近匹配的方法求解了钝锥边界层的稳定性方程 ,得到了钝锥边界层转捩数据 .该方法可提高计算精度 。

高超声速尖锥边界层转捩数值模拟

边界层转捩对高超声速飞行器气动力和热产生重要影响.通过联立求解Favré平均层流脉动能方程与Favré平均Navier-Stokes方程,开展了高马赫数条件下尖锥边界层转捩位置的数值预测研究.将数值计算结果与可用的实验结果进行对比,结果表明增大单位Reynolds数可使转捩提前发生,攻角效应可使迎风面转捩延迟,背风面转捩提前.

高超声速圆锥边界层转捩数值研究

以类似于处理湍流脉动的方式来模型化非湍流脉动,考虑湍流的间歇性特征及引起超声速流动失稳机制,研究包括转捩起始位置及转捩区范围的高超声速流动转捩问题,发展可反映扰动模态和可压缩性影响的两方程SST转捩/湍流模式。针对尖锥模型,在马赫数6、攻角0°～2°时,采用第二模态(Second-mode)机制,预测并比较完全层流态、完全湍流态,以及自然转捩发展过程中的表面热流、表面摩阻等气动特性。

马赫数为6的高超声速钝锥湍流边界层空间演化的直接数值模拟

首先求解了来流马赫数为6的零攻角高超声速钝锥边界层的层流基本流场,在选定的计算域入口引入一组有限幅值的T-S波扰动,用高精度差分格式对流动进行了直接数值模拟。引入的扰动触发了转捩,从而得到了空间模式下的湍流边界层。研究了湍流平均场与脉动场的统计特性,给出了相干结构的流动显示图,并对强雷诺比拟的结论进行了检验。

钝锥高超音速绕流脉冲扰动研究

为了探究来流脉冲扰动波对高超音速流动的干扰,采用7阶迎风WENO格式对无粘通量项离散和6阶中心差分格式对粘性项离散,并使用3阶总变差减少龙格-库塔法进行时间推进求解N-S方程.基于6马赫绕钝锥自由来流,在稳定流场来流外边界加入一个脉冲扰动波,直接数值模拟了脉冲扰动的演变发展过程,并基于ABAQUS有限元软件模拟了钝锥体的瞬态压力响应,分析了对气动热力学特性及钝锥力学特性的影响.结果显示,数值方法精确捕捉到了脉冲扰动的演变发展,在脉冲波作用下,弓形激波出现变形,来流扰动波参数被放大数倍甚至数十倍,近壁面区域产生了涡;流场气动热力学特性和结构压力响应显著地受到脉冲波扰动影响,钝锥头部区域表面的热力学特性变化趋势显著地异于非头部区域,且前者变化较后者显著,后者沿驻点线大致呈线性增长趋势.

钝锥动态转捩风洞试验

钝锥飞行器俯仰振动时,非对称边界层转捩区前后移动相对于俯仰角运动的滞后将产生显著的非定常气动效应,导致飞行器俯仰动稳定性降低,严重时甚至会导致运动失稳。将基于气浮轴承的动态试验技术与转捩红外测量技术相结合,建立了钝锥动态转捩风洞试验技术。该技术采用气浮轴承提供俯仰运动自由且阻尼极低的支撑环境,采用红外热像仪实时测量飞行器表面整体转捩状态。开展了9°钝锥标模动态转捩试验,研究模型绕俯仰轴转动时边界层转捩区前后移动的非定常气动作用与飞行器运动的耦合效应。测量试验观察到了动态转捩试验迎角振荡现象,并发现了转捩区前后移动相对于迎角振荡的滞后现象,获得了转捩滞后时间。