激波风洞边界层转捩测量技术及应用

高超声速边界层转捩对摩阻、传热等有重要影响。在高超声速飞行器研制中，迫切希望能精确预测和控制边界层转捩。激波风洞作为高超声速气动热环境试验的主要地面模拟设备，是研究高超声速边界层转捩的重要设备。但激波风洞原有测量技术适用于工程型号试验，需要依据高超声速边界层转捩特点进行适应性改造和升级。依据高超声速边界层转捩过程中的热流、压力、密度等物理参数变化，发展了薄膜热流传感器测热技术、温敏热图测量技术、高频脉动压力测量技术、高清晰度纹影显示技术等适用于激波风洞的边界层转捩测量技术。并针对头部钝度0.05 mm的半锥角7°尖锥模型，在中国空气动力研究与发展中心Ø2 m激波风洞（FD-14A）马赫数10、单位雷诺数1.2×10⁷/m的流场条件下开展了边界层转捩试验。采用多种转捩测量技术同时进行测量，获得尖锥模型表面边界层转捩情况、边界层脉动压力频谱特征、边界层内清晰的第2模态波和湍流斑纹影图像，不同测量技术获取的试验结果可相互印证，线性稳定性理论分析结果与试验结果相吻合。     

高超声速尖锥边界层压力脉动和热流脉动特性试验

高频脉动热流是激波风洞研究高超声速边界层转捩的重要测试量,利用尖锥模型在中国空气动力研究与发展中心?2 m激波风洞(FD-14A)内开展来流马赫数10、单位雷诺数分别为1.2×107/m、4.7×106/m、2.4×106/m流场条件下的风洞试验.获得了不同工况和流态条件下尖锥模型边界层热流脉动和压力脉动频谱特性,通过...     

原子层热电堆热流传感器在激波风洞试验中的应用

高超声速边界层转捩及激波-边界层干扰等基础气动问题的风洞试验中,高频脉动热流是一重要测试量.为了验证自研的原子层热电堆热流传感器在实际风洞试验环境中的性能,利用尖锥模型在Φ2m激波风洞(FD-14A)内开展来流马赫数10、单位雷诺数4.7×106/m等流场条件下的风洞试验,并与高频脉动压力传感器对比,热流测试结果反映出...     

分布式粗糙元对高超声速边界层不稳定性的影响试验

高超声速边界层转捩是空气动力学亟须研究的关键问题之一。飞行器表面的热防护系统会由于高温烧蚀产生宏观分布式粗糙元形貌,但是目前该宏观烧蚀形貌对高超声速边界层内不稳定波演化机制的影响尚不明晰。基于华中科技大学?0.5 m马赫数6 Ludwieg管风洞,利用高频压力传感器（PCB）和红外热像仪等测量手段,重点研究了分布式粗糙元布置位置和宽度对零攻角下7°半张角尖锥模型高超声速边界层不稳定波演化特征和转捩位置的影响。试验结果表明,将分布式粗糙元布置于同步点之前会促进第二模态不稳定波的演化,且越靠近前缘位置粗糙元宽度因素对下游不稳定波演化影响越小;但随其分布位置向下游移动,促进转捩作用减弱,对不稳定波的非线性交互抑制效果更为明显,转捩位置随之后移。     

高超声速圆锥边界层失稳条纹结构实验研究

边界层转捩的准确预测是高超声速飞行面临的关键气动问题之一。为研究高超声速边界层失稳和转捩机理,以前缘半径1.6mm、半锥角7°的圆锥模型为研究对象,在FD-07高超声速风洞中采用红外热图技术开展边界层转捩实验测量。通过与工程计算结果对比,确认模型表面边界层流态。实验结果表明:有迎角条件下,模型表面中后段出现条纹结构,条纹结构的起始位置随着周向角的增加而向上游移动;随着迎角的增加,条纹起始位置向上游移动,条纹强度差异和条纹与模型中心线的夹角越来越大。实验获得的条纹结构与不同频率扰动波相互作用直接数值模拟获得的条纹结构现象一致。通过对比分析,认为边界层内不同频率扰动波相互作用是产生条纹结构的一种机制。     

高超声速边界层基频二次失稳条纹结构的稳定性

近年来在高超声速边界层的直接数值模拟和静风洞实验研究中,相继发现了边界层转捩前出现的典型基频模态二次失稳现象,其主要成分为流向条纹结构.全文以高超声速平板边界层为研究对象,采用线性稳定性分析和二次稳定性分析的方法,对边界层内条纹结构的产生机制和无黏稳定性特征进行了研究.结果表明:首次失稳扰动幅值对二次失稳类型有影响.当首次失稳扰动幅值较大时,基频模态占主导,其主要成分为条纹结构,表现为流向涡.该条纹结构存在着多个无黏失稳模态,其中低频模态对应于第一模态在三维边界层内的扩展,高频模态对应于可压缩的第二模态.这一研究成果为进一步开展高超声速边界层转捩机制研究奠定了基础.     

超声速/高超声速边界层转捩后期流场的模态分析

将动力模态分解(DMD)方法应用到超声速和高超声速边界层转捩后期的流场分析中,通过获得流场主要的相干结构和对模态的重构,研究了相干结构与壁面阻力和热流的关系。结果表明超声速和高超声速边界层转捩的流场结构存在明显差别。超声速转捩流场由低频流向涡的模态主导,这些模态对转捩后期的壁面阻力和热流有重要贡献;高超声速转捩流场中存在多个不同量级频率的模态,在DMD频谱上表现为多个不同的分支,通过对不同分支能量最高的模态进行考察,我们发现低频模态的结构为流向条带,高频模态的结构为二维扰动波,这些模态对壁面阻力和热流的影响与模态的结构形式类似。     

高超声速边界层流动稳定性实验研究

高超声速条件下边界层转捩相关研究是近年来空气动力学领域的研究热点。通过采用基于纳米粒子示踪平面激光散射(Nano-tracer-based planar laser scattering,NPLS)技术以及温敏漆等测试技术,对高超声速边界层的流动稳定性开展了实验研究。通过NPLS技术,对圆锥边界层中第二模态波精细结构进行了测量,并基于时间相关的测量结果,对第二模态波的波长和频率进行了分析。针对裙锥边界层NPLS结果,计算了特定位置上功率谱空间分布结果,测量得到了高次谐波成分。通过温敏漆和NPLS结果,发现主导三角翼前缘边界层转捩的模态为行进横流模态,分析了该模态的特性,并且与Kulite传感器测量得到的频率进行比较。     

超声波吸收层在天地往返运载器热防护中的应用与研究

超声波吸收层能够抑制高超声速边界层中的第二模态波,延迟边界层转捩,从而降低天地往返运载器热防护系统的重量、成本和复杂度。为了研究超声波吸收层的吸波特性,建立了线性稳定性理论模型。首先,基于四阶龙格-库塔法,采用打靶方式,获得了高超声速边界层基本流。然后,将超声波吸收层的导纳引入稳定性方程边界条件,利用切比雪夫谱方法,求解时间模式下第二模态波的增长率。对于槽道构型超声波吸收层,模型预测结果与已有数据进行了对比,二者符合良好。在几何参数相同的情况下,圆孔构型超声波吸收层对应的第二模态波增长率整体上低于槽道构型。当微腔深度为0.3,微腔数量为6,两种构型超声波吸收层对应的第二模态波增长率具有最小值。第二模态波增长率的最小值出现在微腔数量最少的超声波吸收层上,这是由第二模态波在超声波吸收层顶部散射产生了新的高频扰动波所造成。     

热化学非平衡高超声速平板边界层线性稳定性分析

高超声速边界层流动转捩是近期空气动力学研究的热点问题。对于环境扰动较小的自然转捩过程,稳定性分析已被证明是研究扰动演化的重要手段。另一方面,高超声速边界层内的温度会随着马赫数的升高而快速上升,极高的温度会引起所谓的高温真实气体效应,使得量热完全气体假设失效,从而对边界层稳定性和转捩产生影响。本文针对高温热化学非平衡气体,利用空气5组分模型开展了平板边界层的线性稳定性分析,重点研究了热化学过程对模态稳定性的影响,并探究了边界层离散谱模态的演化和同步过程。研究表明,对于由第二模态主导的高超声速二维边界层:(1)扰动相比基本流更趋向于热化学冻结态;(2)扰动方程中新出现的非平衡源项的扰动项对稳定性影响很小,非平衡过程主要是通过改变基本流剖面来间接影响稳定性;(3)声速是影响第二及更高模态的重要参数,热化学平衡态假设引起的声速计算式的变化能够解释边界层温度和厚度降低时第二模态频率反而降低的非常规趋势。     

等离子体激励器控制平板边界层转捩实验研究

在低速射流风洞中,研究了单级介质阻挡放电等离子体激励器对光滑平板边界层转捩位置的控制作用。实验采用热线测量技术,以边界层速度脉动与平均速度型作为转捩判据。实验发现,在来流速度为15 m/s,激励器连续放电参数为输出电压峰峰值11 kV,频率4.7 kHz时,在激励器放电作用下,平板边界层转捩位置推迟约40 mm。在相同的来流条件和激励器布局下,研究了不同放电参数对边界层内速度型,速度脉动以及频谱分布的影响,发现提高放电电压、频率和占空比能进一步推迟转捩。实验结果表明:激励器产生的射流效应可以增强边界层流动的稳定性,随放电电压、频率以及占空比增强,射流能量增大,因此边界层稳定性进一步加强,转捩控制效果也更明显。     

单模块超燃发动机推力测量天平研制

介绍了用于单模块超燃发动机推力测量的力传感器单分量天平的研制、应用,通过脉冲燃烧风洞和长时间风洞试验,天平/模型/支撑系统频率为20 Hz,在脉冲燃烧风洞有效试验时间内(约300 ms)可获得较为稳定的测力信号,基本满足脉冲燃烧风洞单模块发动机测力要求.在相同试验条件下,脉冲燃烧风洞和长时间风洞获得了相同的发动机推力收益,验证了天平测量数据的准确性.      

Experimental investigation on aero-heating of rudder shaft within laminar/turbulent hypersonic boundary layers

The aero-heating of the rudder shaft region of a hypersonic vehicle is very harsh, as the peak heat flux in this region can be even higher than that at the stagnation point. Therefore, studying the aero-heating of the rudder shaft is of great significance for designing the thermal protection system of the hypersonic vehicle. In the wind tunnel test of the aero-heating effect, we find that with the increase of the angle of attack of the lifting body model, the increasement of the heat flux of the rudder shaft is larger under laminar flow conditions than that under turbulent flow conditions. To understand this, we design a wind tunnel experiment to study the effect of laminar/turbulent hypersonic boundary layers on the heat flux of the rudder shaft under the same wind tunnel freestream conditions. The experiment is carried out in the ?2 m shock tunnel(FD-14 A) affiliated to the China Aerodynamics Research and Development Center(CARDC). The laminar boundary layer on the model is triggered to a turbulent one by using vortex generators, which are 2 mm-high diamonds. The aero-heating of the rudder shaft(with the rudder) and the protuberance(without the rudder) are studied in both hypersonic laminar and turbulent boundary layers under the same freestream condition. The nominal Mach numbers are 10 and 12, and the unit Reynolds numbers are2.4 × 10~6 m~(-1) and 2.1 × 10~6 m-1. The angle of attack of the model is 20°, and the deflection angle of the rudder and the protuberance is 10°. The heat flux on the model surface is measured by thin film heat flux sensors, and the heat flux distribution along the center line of the lifting body model suggests that forced transition is achieved in the upstream of the rudder. The test results of the rudder shaft and the protuberance show that the heat flux of the rudder shaft is lower in the turbulent flow than that in the laminar flow, but the heat flux of the protuberance is the other way around,i.e., lower in the laminar flow than in the turbulent flow. The wind tunnel test results is also validated by numerical simulations. Our analysis suggests that this phenomenon is due to the difference of boundary layer velocities caused by different thickness of boundary layer between laminar and turbulent flows, as well as the restricted flow within the rudder gap. When the turbulent boundary layer is more than three times thicker than that of the laminar boundary layer, the heat flux of the rudder shaft under the laminar flow condition is higher than that under the turbulent flow condition. Discovery of this phenomenon has great importance for guiding the design of the thermal protection system for the rudder shaft of hypersonic vehicles.     

(高)超声速流动试验技术及研究进展

近年来,与高速飞行器相关的(高)超声速流动受到了极大的关注。这类流动所具有的非定常性、强梯度和可压缩性对试验方法和风洞设计技术提出了挑战。超声速纳米示踪平面激光散射(NPLS)技术是由作者所在团队研发的非接触光学测试技术。它能够以较高的空间分辨率来揭示超声速三维流场的一个瞬态剖面的时间解析的流动结构。介绍了NPLS技术以及基于NPLS开发的密度场测量、雷诺应力测量和气动光学波前测量等方法,并回顾了这些技术在超声速边界层、超声速混合层、超声速压缩拐角、激波/边界层相互作用和光学头罩绕流等流动中的应用,清晰地再现了边界层、混合层、激波等典型流场结构及其时空演化特性。另外,为了模拟和研究高空大气条件下边界层自然转捩和超声速混合层的转捩特性,介绍了高超声速静风洞、超-超混合层风洞的设计技术以及层流化喷管的设计方法。     

带减阻杆高超声速飞行器外形气动特性研究

采用高超声速风洞测力试验方法测量钝头飞行器头部减阻杆的高超声速气动特性,研究减阻杆的气动减阻原理,分析了多组不同构型减阻杆的减阻效果。结果表明,减阻杆显著减少了钝头飞行器高超声速的阻力,最大的减阻率达到60％之多;减阻效果与减阻杆构型和迎角状态密切相关;减阻杆会诱发稳定性、“热斑”以及非定常脉动等不利问题。     

Advances in critical technologies for hypersonic and high-enthalpy wind tunnel

Hypersonic and high-enthalpy wind tunnels and their measurement techniques are the cornerstone of the hypersonic flight era that is a dream for human beings to fly faster, higher and further. The great progress has been achieved during the recent years and their critical technologies are still in an urgent need for further development. There are at least four kinds of hypersonic and high-enthalpy wind tunnels that are widely applied over the world and can be classified according to their operation modes. These wind tunnels are named as air-directly-heated hypersonic wind tunnel, light-gas-heated shock tunnel, free-piston-driven shock tunnel and detonation-driven shock tunnel, respectively. The critical technologies for developing the wind tunnels are introduced in this paper, and their merits and weakness are discussed based on wind tunnel performance evaluation. Measurement techniques especially developed for high-enthalpy flows are a part of the hypersonic wind tunnel technology because the flow is a chemically reacting gas motion and its diagnosis needs specially designed instruments. Three kinds of the measurement techniques considered to be of primary importance are introduced here, including the heat flux sensor, the aerodynamic balance, and optical diagnosis techniques. The techniques are developed usually for conventional wind tunnels, but further improved for hypersonic and high-enthalpy tunnels. The hypersonic ground test facilities have provided us with most of valuable experimental data on high-enthalpy flows and will play a more important role in hypersonic research area in the future. Therefore, several prospects for developing hypersonic and high-enthalpy wind tunnels are presented from our point of view.     

Theories and methods for designing hypersonic high-enthalpy flow nozzles

Hypersonic high-enthalpy wind tunnels have been a challenge to ground tests in aerospace research area for decades and its test flow uniformity is one of the most important parameters for evaluating wind tunnel performances. Regarding to the performance requirement, theories and methods for designing hypersonic flow nozzles at high enthalpy conditions are quite difficult, but very interesting topics, especially when air molecule dissociations take place in wind tunnel test gas reservoirs. In thi...