Laminarization of a boundary layer on the perforated surface with blind damping cavities in an accelerating flow

In this paper, a decrease of turbulent pulsations of velocity in a boundary layer and the coefficient of friction drag for an accelerating flow on the perforated surface with blind damping cavities is experimentally found. We generalize the mathematical model of partial boundary layer laminarization [1], which is based on the experimental data [2, 3] obtained earlier on a decrease of the friction drag coefficient and deformation of average velocity profiles in the stabilized section of flow in the perforated tube with blind damping cavities.     

Supersonic drag reduction technology in the scaled supersonic experimental airplane project by JAXA

Anticipating the international cooperative development of a next generation supersonic transport (SST), Japan Aerospace Exploration Agency (JAXA) has developed an advanced drag reduction technique as one of the key technologies that will be required. JAXA's technique is based on an aerodynamically optimum combination of well-known pressure drag reduction concepts and a new friction drag reduction concept. The pressure drag reduction concepts are mainly grounded in supersonic linear theory and involve the application of an arrow planform, a warped wing with optimum camber and twist, and an area-ruled body. The friction drag reduction concept is a world-first technical approach that obtains a natural laminar flow wing with a subsonic leading edge at supersonic speed. An ideal pressure distribution is first designed to delay boundary layer transition even on a highly swept wing, then an original CFD-based inverse design method is applied to obtain a wing shape that realizes the pressure distribution. An unmanned and scaled supersonic experimental airplane was flown at the Woomera test field in Australia in October 2005 to prove those concepts. Flight data analysis and comparison of flight data with CFD design data validated the drag reduction technique both qualitatively and quantitatively.     

超临界层流机翼边界层及气动特性分析

高空长航时无人机设计巡航状态的雷诺数较小,黏性边界层对气动特性的影响较大。详细分析了雷诺数对机翼边界层和气动力的影响,用数值方法对超临界层流机翼三维层流-转捩-湍流混合边界层特性进行了研究,分析比较了高空小雷诺数和中空大雷诺数情况下机翼三维边界层的特性,尤其是边界层转捩点位置、表面摩阻和气动特性的雷诺数效应。研究表明雷诺数对于高空无人机机翼边界层厚度、摩擦阻力和升阻比影响较大;对层流机翼的转捩点位置和升力系数影响较小;自然层流机翼技术可以应用于高空无人机设计。     

壁面温度控制对平板边界层影响的数值研究

通过对零压力梯度的平板边界层流动施加温度控制,展开壁面温度控制对平板层流边界层和湍流边界层影响的研究,探索温度控制对平板转捩雷诺数和壁面摩擦阻力的影响规律。采用带有转捩模式的三方程湍流模型对平板边界层流动进行数值模拟,重点考察了壁面摩阻系数、平板转捩雷诺数、湍流边界层流动随壁面温度变化的规律。计算结果表明在壁面温度从288 K 增大到432 K 时,边界层转捩雷诺数增大约36%,表面摩擦阻力减少约9.6%。研究分析表明：加热控制使层流区域温度边界层内粘性作用增强,雷诺切应力和湍动能减小,流动更加稳定;而湍流区域边界层内粘性底层中速度梯度和粘性切应力减小,导致壁面处摩擦切应力减小。因此壁面加热控制可以延迟边界层转捩,减小湍流区摩阻系数,并减小平板摩擦阻力。     

超临界自然层流机翼设计及基于TSP技术的边界层转捩风洞试验

为了实现绿色航空节能减排的目标，层流设计技术成为飞行器设计者的研究热点。对于跨声速客机而言，超临界自然层流机翼设计技术将显著减小飞行阻力，提升气动性能，减少燃油消耗和污染物排放。首先，基于高精度边界层转捩预测技术耦合翼型优化设计系统，实现超临界自然层流翼型设计；经过合理的翼型配置，形成超临界自然层流机翼。转捩数值模拟分析结果表明，超临界自然层流机翼的层流流动特性良好。然后，以比例为1：10.4的试验模型在荷兰高速低湍流度风洞进行边界层转捩风洞试验，使用温度敏感材料涂层（TSP）技术拍照获得机翼表面在不同马赫数、雷诺数和迎角工况下的层流-湍流分布。最后，通过超临界自然层流机翼边界层转捩试验结果，探讨了该类型机翼的转捩特性随来流参数的变化规律，总结了超临界自然层流机翼设计的关键因素。此外，该模型也可用来验证边界层转捩预测技术在超临界、高雷诺数工况下的预测精度。     

An investigation of surfactant and enzyme formulations for the alleviation of insect contamination on Hybrid Laminar Flow Control (HLFC) surfaces

Hybrid Laminar Flow Control (HLFC) is an active drag reduction technique that requires a small amount of air to be sucked through a porous skin surface, thus stabilising the boundary layer and permitting extended laminar flow along the wing surface. Contamination of the laminar flow surfaces by insects is a major concern for this technology. An overview of insect contamination and mitigating solutions is provided as background to research undertaken. The paper reports on an experimental investigation into the potential use of surfactants and enzymes as additives to contamination alleviation fluids. It was demonstrated that both approaches have merit. It is concluded that the use of fluorosurfactant containing solutions is preferable to enzyme borne solutions, partly due to the need to control the temperature of the enzyme solution.     

层流流动主/被动控制技术

摩擦阻力在民机总阻力中占很大比重,减少摩擦阻力对改善民机性能和实现绿色航空具有重要意义。层流摩擦阻力远小于湍流摩擦阻力,因此扩大层流区,甚至实现全层流流动,是减阻的一个重要途径。目前形成了自然层流流动、全层流流动和混合层流流动控制（HLFC）3种层流流动主动控制技术。本文基于减阻和流动不稳定分析,对3种控制技术的概念、方法、优缺点、可带来的效益和应用层流流动控制技术的飞机的设计方法及维护（包括预防昆虫和冰粒等污染的措施）等方面作了较为系统的阐述。概要地介绍了X21A、Jetstar、Boeing 757等飞机的HLFC飞行试验验证项目,结果表明了层流流动主动控制技术的有效性和困难性。本文也从原理到飞行试验较为系统地介绍了一种层流流动被动控制技术。     

Shock control on a swept wing

Shock and boundary layer control by contour bumps and local boundary layer suction have been investigated experimentally and numerically on a transonic swept wing. Additional 2-D numerical investigations were performed for the airfoil, corresponding to the wing. The investigations were primarily stimulated by the question concerned with the influence of sweep on the bump effectiveness. This influence has been found to be rather small; the drag reduction by the bump is slightly lower for the swept wing than for the airfoil. A location of the bump in the shock region has shown its effectiveness for reducing shock strength and hence wave drag. A position of the bump downstream of the shock wave has been shown to reduce viscous drag and to postpone buffet-onset to higher lift coefficients. Furthermore, the results indicate that boundary layer suction is a powerful device for drag reduction, but the effectiveness decreases with increasing Reynolds number. Higher effectiveness of suction can be attained, when it is coupled with a contour bump. The parameters height and position (relative to the shock) of the bump, optimized in terms of drag, depend on the shock strength; an influence of the boundary layer thickness upstream of the shock on the optimal bump parameters has not been found. A possibility to control an adaptive bump, mounted on an aircraft wing, is to employ the trailing edge pressure.     

HLFC后掠翼优化设计的若干问题

使用扩展自由变形参数化方法,基于径向基函数的动网格技术和改进的混合粒子群算法,考虑吸气的e^N转捩预测方法和雷诺平均Navier-Stokes求解器,搭建了针对混合层流流动控制（HLFC）后掠翼的优化设计平台,对HLFC后掠翼的气动外形设计、雷诺数影响、吸气分布设计等多个问题进行了研究,对比分析了在这些因素影响下HLFC后掠翼的阻力系数和层流区长度的差别,进而探索了相应的设计准则。研究表明,对于层流区较长和阻力系数较小的HLFC后掠翼来说,它们上表面的压力分布具有共同的特征：头部峰值较低,之后有一个小的逆压,接下来是一段较长的均匀稳定的顺压,这段顺压最后终结于一道激波。应用HLFC技术后,通过实现大面积的层流区,机翼的摩擦阻力和压差阻力均可显著地降低,降低的幅度远大于不考虑层流控制的设计结果。同时,HLFC机翼的设计应综合考虑摩擦阻力、压差阻力、激波强度和配平阻力（低头力矩）,层流区最长不一定意味着阻力最小。一般来说,雷诺数越高,越难维持层流,但应用混合层流控制技术后,即使在难以实现自然层流的高雷诺数下,HLFC机翼依然有较长的层流区。通过对吸气分布的设计进行研究,说明了非均匀吸气比均匀吸气要更有效率一些,能够节省吸气量。     

二维翼型混合层流控制减阻技术试验研究

选择NACA0006系列层流翼型作为物理模型,使用FLUENT商用软件计算分析翼型表面压力梯度,结合对翼型后缘做局部优化修形增大顺压梯度范围以及在翼型前缘布置吸气控制单元并配套吸气装置形成混合层流控制减阻技术。风洞试验中应用红外成像技术测量翼型表面层流区域,探索研究了混合层流控制减阻技术的实用效果。试验结果表明：对翼型实施混合层流控制减阻技术后,明显增大了翼型表面的层流面积。     

微吹气对湍流平板边界层流动特性的影响及其减阻机理

微吹气技术能够改变平板湍流流场结构,减小平板表面的摩擦阻力。采用直接数值模拟方法,计算了来流马赫数0.7条件下,流场流过光滑平板和NASA-PN2多孔平板表面两种情况,通过对比这两个算例的相关流场特征,验证了微吹气控制减阻的有效性,局部最大减阻率达到了45%,并且由于微吹气控制的"记忆"功能,减阻效果在微吹气流域下游仍会持续一段距离,增加了减阻区域的流向面积。壁湍流摩擦减阻的原因在于近壁区域出现了一个低速的"湍流斑",黏性底层厚度增加,速度型曲线被抬升。但与此同时,边界层内湍流速度脉动也得到了增强。进一步对流向脉动涡演化规律分析,发现微吹气对流向脉动涡发挥着多重作用。在增加流向脉动涡强度的同时,还使得流向涡团向远离壁面抬升,这样减小了流向涡与壁面之间直接作用。此外,微吹射流产生的冲击作用会在流向涡表面留下凹痕,使得流向涡分散成相对小的涡团结构。     

湍流减阻新概念的实验探索

本文对当今湍流表面摩擦减阻新概念进行了初步的风洞实验探索。将垂直于流动方向的小尺寸肋条按一定的间隔距离固定在平板上,利用自制的悬挂式天平测量了不同风速时的阻力,获得了约１０．２％的减阻效果。实验中分别考察了肋条参数对减阻的影响,使用Ｘ型热线风速仪研究了雷诺应力的型态。从湍流边界层涡结构的观点出发,提出了边界层底部“微型空气轴承（ＭＡＢＳ）”减阻新概念以及涡结构干扰对减阻的影响,并认为平均速度型态的     

后掠翼三维鼓包串激波控制参数的影响

针对跨声速后掠翼,三维鼓包串作为一种有效的减阻方式具有结构简单、高效及鲁棒性好等优点.利用全局优化算法探索了鼓包设计参数空间的整体特性,并对鼓包长度、三维鼓包展向设计参数对鼓包减阻效果的影响进行了研究,发现鼓包顶点位置和高度对阻力系数最敏感,三维鼓包的展向设计参数则对阻力系数不敏感,而鼓包长度和鼓包相对展长越长越有利于减阻.在此基础上开展了小后掠角自然层流机翼加3种不同类型鼓包串的优化研究,通过优化结果发现,增加优化后的三维鼓包串,可将小后掠角自然层流机翼阻力发散马赫数向后推移,并且鼓包平均长度和控制区越大,效果越好.三维鼓包串具有良好的局部控制特性,可用于局部较强激波的抑制.三维鼓包串对常规后掠翼波阻具有良好的控制效果,同时能够抑制激波诱导的机翼后缘气流分离.      

高雷诺数下机翼表面层流段长度对减阻量的影响

层流减阻技术是提高飞机经济性的重要手段,开展可应用于工程实践的层流减阻研究具有重要的意义。针对某民用飞机翼身组合体构型,采用数值模拟法分别研究雷诺数为1.0×107和1.8×107时全湍流以及机翼弦向保持7%、15%、20%、30%、40%层流段长度范围的减阻特性。结果表明:与全湍流情况相比,层流段长度的增加可以有效减小飞机阻力,增加升阻比;当层流段长度保持在40%时,飞机的减阻量可以达到11.0%左右,而升阻比可增加12.3%左右,且在较小雷诺数下有着更大的减阻收益;层流范围增加可有效减小摩擦阻力系数。     

层流流动控制技术及应用

民机受到的摩阻占其总阻力很大比例,减少摩阻对改善民机性能和降低成本具有重要意义.层流摩阻远小于湍流摩阻,因此扩大层流区,甚至实现全层流流动是减阻的一个重要途径.为此,形成了包括自然层流流动、全层流流动和混合层流流动3种层流流动控制技术.本文从减阻分析,对3种层流流动控制技术的概念、方法、优缺点、可带来的效益,层流流动控...     

基于层流机翼的增升装置设计研究进展

层流机翼的增升装置设计可以起到减阻和降噪的作用。将低阻的层流区域与高效的增升装置相结合,通过采用有效的外形设计和恰当的结构选型,可以实现巡航时层流区域的扩大、起降时增升效果的提升,并显著降低机体噪声。本文根据绿色航空的发展需求,分析了层流机翼和增升装置的技术特征,阐述了在层流机翼状态下增升装置的设计方法,系统地介绍了结构外形、运动机构和辅助装置等的应用现状,在此基础上综述了实现增升装置适应性调整所需要重点研究的技术内容,给出了层流机翼增升装置的未来发展方向和思路。     

Instability analysis and drag coefficient prediction on a swept RAE2822 wing with constant lift coefficient

Swept wing is widely used in civil aircraft,whose airfoil is chosen,designed and optimized to increase the cruise speed and decrease the drag coefficient.The parameters of swept wing,such as sweep angle and angle of attack,are determined according to the cruise lift coefficient requirement,and the drag coefficient is expected to be predicted accurately,which involves the instability characteristics and transition position of the flow.The pressure coefficient of the RAE2822 wing with given constant lift coefficient is obtained by solving the three-dimensional Navier-Stokes equation numerically,and then the mean flow is calculated by solving the boundary layer(BL) equation with spectral method.The cross-flow instability characteristic of boundary layer of swept wing in the windward and leeward is analyzed by linear stability theory(LST),and the transition position is predicted by eNmethod.The drag coefficient is numerically predicted by introducing a laminar/turbulent indicator.A simple approach to calculate the lift coefficient of swept wing is proposed.It is found that there is a quantitative relationship between the angle of attack and sweep angle when the lift coefficient keeps constant;when the angle of attack is small,the flow on the leeward of the wing is stable.when the angle of attack is larger than 3°,the flow becomes unstable quickly;with the increase of sweep angle or angle of attack the disturbance on the windward becomes more unstable,leading to the moving forward of the transition position to the leading edge of the wing;the drag coefficient has two significant jumping growth due to the successive occurrence of transition in the windward and the leeward;the optimal range of sweep angle for civil aircraft is suggested.     

吹吸扰动对壁湍流边界层摩擦阻力的影响

对风洞中平板湍流边界层施加局部周期性吹吸扰动,并用IFA-300热线风速仪测量了其近壁区域不同法向位置瞬时速度的时间序列信号.利用湍流边界层对数律平均速度剖面,计算湍流的壁面摩擦阻力.对平板湍流脉动速度信号用子波分析进行多尺度分解,通过条件采样以及相位平均的方法获得不同尺度下的相干结构波形和能量分布,来研究吹吸扰动对平板湍流边界层的摩擦阻力的影响.结果表明,扰动使得窄缝附近流动阻力增加,而下游区域流动阻力减小.      

沟槽表面边界层湍动能分布规律

 通过对不同风速下不同尺寸V型沟槽表面及光洁平板表面边界层内湍动能的测试,对比分析了沟槽表面边界层湍动能的分布规律。试验在一小型专用风洞中开展,流场测试中使用恒温式IFA300智能型流动分析仪,测试模型则采用有机玻璃材质的矩形平板结构;而在沟槽表面理论零点及壁面摩擦速度的计算中,采用基于湍流边界层Spalding壁面律公式同时计算壁面理论零点和壁面摩擦速度的改进方法。最终研究结果表明,沟槽结构主要影响边界层流场的近壁区,沟槽的存在提高了边界层中黏性底层内湍流脉动所具有的动能,有效降低了边界层中过渡区内的湍动能,最大相对降幅超过10%,较好地验证了基于“二次涡”的沟槽表面减阻理论。     

基于狭缝合成射流的湍流边界层流动控制实验研究

湍流边界层主动流动控制减阻是当下一个热门的研究领域。为了探究狭缝合成射流对平板湍流边界 层流动控制的减阻效果,分别在有无合成射流干扰的条件下利用恒温式热线风速仪测量湍流边界层的流向速 度,根据测量结果得到时均速度型、脉动速度型、偏斜因子、平坦因子,并进行对比分析;选取不同频率和强度组 合的典型合成射流激励条件,研究激励条件带来的减阻效果沿流向的变化。结果表明:受控湍流边界层减阻效 果、施加的合成射流特性和当地站位到狭缝的距离均有关;合成射流在靠近狭缝位置引起湍流边界层表面摩擦 阻力增加;远离狭缝,合成射流产生减阻效果,随流向距离增加,减阻效果先增强后减弱;高频合成射流比低频 合成射流减阻效果更强;脉动速度的功率谱密度和自相关性分析表明合成射流的作用效果沿流向逐渐减弱。