RESEARCH PROGRESS OF PLASMA/MHD FLOW CONTROL IN INLET 1)
Li Yiwen*,2), Wang Yutian*, Pang Lei??, Xiao Lianghua*, Ding Zhiwen*, Duan Pengzhen*收稿日期:2018-09-30接受日期:2019-01-25网络出版日期:2019-03-18
| 基金资助: |
Received:2018-09-30Accepted:2019-01-25Online:2019-03-18
作者简介 About authors
2)李益文,博士,主要研究方向：进排气流动控制与隐身.E-mail：lee_yiwen@163.com
摘要
为实现高速飞行器的宽速域飞行,如何保证进气道在非设计状态下的性能至关重要。相比于传统被动控制方式,等离子体/磁流体流动控制技术作为新概念主动流动控制技术,由于其具有结构简单,快速响应,并可根据实际飞行条件进行反馈控制等优势,在国内外上得到了广泛关注。本文介绍了等离子体/磁流体在高超/超声速进气道的主要应用方式与等离子体/磁流体建模方法。当进气道处于超临界状态时,等离子体/磁流体流动控制主要通过热阻塞效应产生虚拟型面,从而将激波系推回至唇口,该技术有望在需要短时间流动控制的高马赫数导弹上走向工程应用;由于等离子体/磁流体激励器与壁面平齐安装,对于高超声速飞行条件,相比于粗糙元其对热防护的要求较低,并且通过超声速风洞实验初步证明了通过高频激励对边界层施加扰动的可行性,需要从稳定性理论的角度对其物理机制进行研究。在后续发展中需要进一步创新等离子体产生技术及激励方式,发展等离子体与流的全耦合计算模型等离子体与流的全耦合计算模型与高效算法 ,为指导工程应用提供依据.
关键词：
Abstract
In order to realize wide-speed-range flight of high-speed vehicle, it is of great importance to maintain the performance of inlet at off-design. Compared with traditional passive control methods, plasma and magnetohydrodynaimic(MHD) flow control are novel active flow control methods, and they have attracted extensive attention worldwide, as a result of some advantages, such as simple structure, fast response and feedback control based on actual flight condition, etc. In this paper, the main applications of plasma and MHD in hyper/supersonic inlet and dynamics models are introduced. When the inlets are in supercritical state, the shockwaves can be push back to cowl as a result of the virtual surface produced by plasma and MHD, which is based on the effect of thermal chocking. This technology is expected to applied on the hypersonic missile if only short-time flow control is required. The plasma and MHD actuators can be mounted flush on the wall, so that its requirement for thermal protection is less than that of roughness at hypersonic flight condition. The applications of high-frequency plasma and MHD actuation to produce disturbances in boundary layer have been validated through supersonic wind tunnel experiment, and the physical mechanism can be interpreted from the point of stability theory. The innovative developments of plasma source technology and the way of actuation, as well as coupled model of plasma and fluid dynamics and efficient algorithms are required in future, which can provide guidance for engineering application.
Keywords：
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本文引用格式
李益文, 王宇天, 庞垒, 肖良华, 丁志文, 段朋振. 进气道等离子体/磁流体流动控制研究进展 1). 力学学报[J], 2019, 51(2): 311-321 DOI:10.6052/0459-1879-18-290
Li Yiwen, Wang Yutian, Pang Lei, Xiao Lianghua, Ding Zhiwen, Duan Pengzhen.
引 言
未来的吸气式高超声速飞行器需要在不同飞行高度、马赫数范围内具有较好的适应能力,如：要求在设计马赫数下具有较好的经济性.为了追求高速突防能力,需要进气道在高于设计马赫数下仍然能工作,并具有较高的总压恢复系数.而为了避免低空、高速飞行时严酷气动力热载荷,在低空、低马赫数下需确保进气道起动,因此,对进气道的性能提出了较高要求.吸气式高超声速进气道在设计上面临诸多挑战,存在激波系调控、边界层转捩、摩擦阻力、激波与边界层相互作用、隔离段流动分离及压力控制等问题[1].目前,吸气式高超声速飞行器大多数采用结构相对简单的定几何进气道设计,且通常选定某一典型工况作为设计点进行激波系的设计,其优点是结构简单、重量轻、可靠性高.当低于设计马赫数时,进气效率下降,流量系数下降、溢流阻力增大,甚至出现进气道不起动的状况,限制了飞行包络的拓展;而在高于设计马赫数时,激波损失增强,激波系打入唇口内壁面,并不断反射,边面层增厚,总压恢复系数降低,发动机推力降低,飞行器总体性能急剧降低.此外,激波/边界层干扰、边界层转捩、边界层分离等增大了流场局部的热载荷、声学载荷和流场畸变,影响燃烧室内油气掺混与燃烧效果,甚至导致进气道不起动、发动机喘振、熄火,严重制约发动机的推进性能.为了提高非设计工况下进气道在性能,需要对进气道的流场进行控制,使之能够有效工作.
等离子体/磁流体流动控制技术,作为一种新概念主动流动控制技术,主要利用气体放电产生等离子体,并在电磁激励的作用下对流场作用,其主要特点是：没有运动部件、响应时间短并且激励频带宽.与传统的机械式调节方式相比,具有结构简单、附加质量小等优点,具有良好的应用前景[2].2002年,《简氏防务周刊》曾将国外进行的等离子体可以大幅度改变飞行器空气动力特性的研究评论为：将期待一场军用和商业飞行器的革命.2005年,美国将等离子体动力学列为未来几十年内保持空军技术领先地位的六大重点发展的基础研究领域之一;我国《国家中长期科学和技术发展规划纲要(2006-2020年)》将高超声速推进系统、磁流体及等离子体动力学等列为"面向国家重大战略需求的基础研究"中的"航空航天重大力学问题".该技术有望拓宽进气道工作马赫数范围,提高进气道总压恢复系数和起动特性等.
本文主要介绍了等离子体/磁流体流动控制技术在高超声速进气道的主要应用方式,包括磁流体激波系调控、边界层磁流体转捩控制、边界层分离磁流体流动控制等,综述了国内外相关研究现状及发展趋势,分析了其应用过程中存在的关键科学与技术问题,并对其发展前景进行了讨论与展望.
1 等离子体/磁流体流动控制基本原理
1.1 等离子体/磁流体概念
等离子体是固体、液体和气体之外的物质第四态,是物质的存在的另一种形态.它是电子、离子、原子、分子或自由基等粒子组成的集合体.对于空气来说,如图1所示,当气体温度超过2500 K时,氧分子发生离解;超过4000 K时,氮分子发生离解;超过9000 K时,氧原子和氮原子发生电离,产生电子和带电离子,形成等离子体.根据其所处的热力学状态,当重粒子温度和电子温度相等时,称之为平衡等离子体,反之称之为非平衡等离子体.由于平衡等离子体对产生的温度非常高,如核聚变反应堆中的等离子体,要求温度超过1亿摄氏度,通常我们能得到的是局部热力学平衡等离子体,如电弧放电等离子体.近年来,随着人工电离技术的进步,特别是非平衡电离技术得到了快速发展,各种新的等离子体产生方法正在发展之中,如高压场致电离,电子束电离,微波电离等,为等离子体技术在航空航天领域应用提供了基础条件.图1
新窗口打开|下载原图ZIP|生成PPT图1空气分子随温度变化发生的化学反应
Fig.1Chemical reaction of air molecule at different temperature
磁流体(magnetohydrodynamics,MHD)是磁流体动力学的通称,它是结合经典流体力学和经典电动力学方法研究导电流体(等离子体、液态金属或电解液等)和磁场(外加或自生)相互作用的学科.另外,磁流体也是一种特殊流体介质的名称,它既是连续介质,又是电磁介质,具有很多普通流体没有的重要特性,现象的多样性和行为的复杂性是磁流体系统的显著特点[3].当前,一些需要用磁流体来认识和解决的科学和技术问题的不断涌现,因而倍受物理学、力学、应用数学和技术科学界的重视.其中,与航空航天应用联系密切、研究最为广泛的是等离子体和电磁场的相互作用.
1.2 基本原理
等离子体/磁流体流动控制是基于等离子体气动激励的新概念主动流动控制技术.等离子体气动激励是利用等离子体在电磁场力作用下运动或气体放电产生的压力、温度变化,对流场施加的一种可控扰动,其物理原理包括3个方面：一是"动力效应",即在流场中电离形成的等离子体或加入的等离子体在电磁场力作用下的定向运动,通过离子与中性气体分子之间的动量输运诱导中性气体分子运动,形成等离子体气动激励;二是"冲击效应",即流场中的部分空气或外加气体电离时产生局部温升和压力升(甚至产生冲击波),形成等离子体气动激励;三是"物性改变",即在流场中的等离子体改变气流的密度、黏性和热传导等热物理特性.由于其没有运动部件,具有结构简单、附加质量小、激励频带宽等优点,因而在航空航天中表现出具有良好的应用前景,受到国内外研究机构和****的广泛研究.早在1957,郭永怀[4]就指出,当飞行器在20 km高空,飞行速度达到Ma=15时,空气便是很好的导体,此时便可以利用电场与磁场进行流动控制.2 进气道等离子体/磁流体激波系调控
目前进气道流动控制大多采用机械变几何流动控制和气动变几何流动控制方法.但是在高马赫数来流的恶劣环境中,变几何调节方式会带来热防护、密封等新的问题,同时变几何调节方式还存在结构复杂、尺寸重量大等不足.对于进气道等离子体/磁流体激波系调控,其目的是增大非设计马赫数下进入进气道气体的流场品质,如增加捕捉空气流量及压缩比等,从而提高非设计马赫数下进气道的性能.与传统的流动控制相比较,采用等离子体/磁流体流动控制可能更好地解决在进气道设计中存在的挑战与困难,可以在没有复杂机械调节装置的情况下,实现对气流的主动流动控制,可大大扩展固定几何发动机的运行范围,提高推进系统性能.2.1 磁流体控制
磁流体激波系调控,主要是基于磁流体流动控制来改变进气道在非设计马赫数下的激波位置,通过平衡或非平衡电离技术产生大体积等离子体,然后通过施加外负载产生定向电流,在磁场的作用下通过$j\times B$洛伦兹体积力加/减速气流,使进气道激波重新交于唇口.俄罗斯艾尔菲物理研究所的Bobashev等最早对大尺寸MHD激波控制进行了相关研究,并基于激波风洞,采用电子复合时间较长的稀有气体氪作为实验工质,超声速气体总温为9800 K以实现平衡电离,理论电导率高达3200 S/m,首次通过实验证明了MHD作用控制进气道激波的可行性.
该实验证明了通过磁流体作用控制进气道斜激波的可行性,但是由于平衡电离所需的温度太高,难以应用于真实飞行环境中的冲压发动机进气道控制.对于高超声速进气道,当飞行马赫数达到12以上才有可能出现平衡电离,即使加入碱金属电离种子也需要飞行马赫数达到8以上才能出现平衡电离等离子体.因此,后续的进气道的磁流体流动控制研究,主要采用非平衡电离如气体放电方式产生等离子体.
Macheret等[8]和Shneider等[9]提出一种基于电子束电离的MHD进气道激波控制方式,其原理示意图如图2所示,通过建立无黏磁流体动力学模型并耦合电子束电离等离子体模型,对该控制方式在非设计马赫数下的控制效果进行了理论分析,结果表明在Ma>Mad时,通过合理的作用参数选择可以增大空气捕获量并减小总压损失,但对于Ma<Mad的情况,由于焦耳热的作用反而使空气捕获量减小,总压损失增大.
图2
新窗口打开|下载原图ZIP|生成PPT图2基于电子束电离的磁流体进气道结构图
Fig.2Schematic of MHD inlet based on e-beam ionization
李益文等[10]采用电容耦合射频放电方式,开展的非平衡电离磁流体流动控制实验研究表明：在负载功率35 W、磁感应强度1.7 T条件下,开展了磁流体流动控制实验研究.由于焦耳热作用显著,加速洛伦兹力作用时,静压上升了130 Pa,马赫数减小了0.38;减速洛伦兹力作用时,静压上升了200 Pa,马赫数减小了0.58.
2.2 等离子体控制
受制于非平衡电离技术与强磁场产生技术的发展,对于大尺寸进气道MHD流动控制,由于其低电导率与低磁场,以及高流动速度将导致作用数较低,控制效果变差.相比较而言,表面等离子体流动控制无需电离大体积气体,而且边界层内气流速度、密度较低,同时壁面处也具有最大的磁场强度,因此具有更好的控制效果.俄罗斯科学院高温研究所Leonov 等[11-14]从2003年开始,对表面放电等离子体气动激励激波控制进行了系统研究,并将其基本原理概括为：热、磁流体、电流体三种作用.利用PWT-10实验系统,研究了低气压Ma=2超声速气流下表面直流放电特性,实验结果表明,在放电功率1~4 kW的条件下,激波位置前移,激波强度减弱,通过高速CCD得到放电细丝的周期性变化规律,提出"准直流"放电模型,并推导出超声速下等离子体型面角度的计算公式,进一步研究发现通过磁场可以有效控制高速气流下的放电细丝的运动并提高放电能量效率;并利用表面等离子体气动激励成功完成了平面诱导斜激波、斜激波位置前移、控制后壁面台阶与凹腔处流动分离等原理性验证实验.
2015年,Leonov与欧洲导弹集团的Falempin等[15]合作开展了基于表面等离子气动激励的二维三激波压缩斜面大尺寸进气道模型流动控制实验研究.进气道模型设计马赫数$Ma_{\rm d}$=2.0,利用英国T-313超声速风洞分别研究了在马赫数2.5和3.0情况下等离子体气动激励的流动控制效果,图3为放电功率6.5 kW激励作用下,来流速度Ma=3条件下有无激励时的激波纹影与马赫数云图,可以看出等离子体层的作用如同等熵压缩面将两激波结构转变为单激波并重新交于唇口, 尽管壁面放电区域静压有所增加,但通过进气道内的测压排管测量表明流场静压分布变化很小,并且可以有效提高质量流率与总压恢复系数,证明了该控制方式在扩展超声速进气道工作范围方面的可行性.
图3
新窗口打开|下载原图ZIP|生成PPT图3有无等离子体激励时的激波结构纹影
Fig.3Schlieren visualization of shockwave structure with or without plasma actuation
普林斯顿大学Miles课题组提出"雪橇(snowplow)"式磁驱动表面放电等离子体激励器[16],并将该激励器用于超声速边界层加速研究,通过高速CCD捕捉到单个等离子柱的运动状态,在2 T磁场的作用,其运动速度由350 m/s增至2000 m/s,边界层马赫数提高约10%[17];进而Kalra等[18]将该激励器用于激波边界层作用(SWBLI)流动控制实验与数值模拟研究,通过高速纹影与平面激光散射等手段证明了该激励方式可以诱导和减缓边界层分离,实验结果如图4所示,与无磁场激励结果对比分析表明$j\times B$洛伦兹力是主要影响因素;同时为进一步研究该激励形式的作用效果,Kalra等[19]又进行了静止条件下诱导速度实验研究,结果表明单次激励下只能产生4 m/s左右的诱导速度.为了更好地理解该物理现象,Macheret对该问题进行了深入的理论分析[20],理论分析结果与实验结果高度吻合,从原理上证明了其可行性;在此基础上,美国空军实验室的Bisek等[21]针对该激励形式开展了LES模拟(大涡模拟)研究,将等离子体气动激励以体积力及能量源项的形式体现在Navier-Stokes方程中,分别模拟了3种情形：(1)体积力的定常激励;(2)体积力的非定常激励;(3)同时考虑体积力和能量源项的定常激励.结果表明,施加激励后的分离区长度得到明显减小(其中方案(1)可达75%),且分离区及再附区下游的湍动能得到了明显抑制,但并未考虑非定常热源项的激励作用、以及非定常激励与激波/边界层干扰的低频扰动频率耦合等问题.国内虽然未见"雪橇"式磁驱动表面放电等离子体激励器的实验研究报道,但也关注了该项研究,并开展了部分数值模拟研究,南京航空航天大学的苏伟仪等[22]开展了该激励方式对超声速平板湍流边界层控制机理的理论研究与数值分析.
图4
新窗口打开|下载原图ZIP|生成PPT图4有无磁驱动等离子体激励时的流场结构纹影
Fig.4Schlieren visualization of flow field structure with or without magnetically driven plasma actuation
3 气道边界层等离子体转捩控制
3.1 需求及基本原理
边界层转捩问题一直是流体力学长期关注但尚未解决的重要研究领域之一,随着飞行器向更高马赫数迈进,边界层转捩问题更加突出[23-26].对于吸气式高超声速飞行器,为了减少进气道前体压缩拐角处和隔离段入口处的流动分离,提高进气道捕获流量,保障超燃冲压发动机的正常工作,进气道前体边界层流动必须是湍流.然而,由于高空来流湍流度很低,吸气式高超声速飞行器在实际飞行中进气道前体边界层流动通常为层流状态,因此,需引入人工转捩技术来促进湍流,以抑制边界层分离.目前,高超声速边界层人工转捩装置可以分为被动和主动两种,被动转捩装置不需要能量输入,一般针对某一设计状态有效,在非设计状态下控制效果降低,主动转捩装置则需要能量的注入.常见的被动转捩装置通常为沿压缩面展向布置的一排或几排具有一定间距的圆柱型、钻石型或斜坡型粗糙单元.由于高度与边界层厚度相当,在促进转捩的同时,会产生局部激波,带来附加阻力和高热流等方面的问题.相关高超声速边界层强制转捩的实验研究工作,国外开展的较早,2004年,X-43A采用斜坡型涡流发生器成功地在飞行试验中实现了强制转捩,随后HIFiRE,X-51等吸气式高超声速飞行器进气道都采用涡流发生器成功实现强制转捩[27].国内清华大学、北京大学、天津大学、北京航空航天大学、中国空气动力研究与发展中心等研究机构开展了大量转捩问题研究.赵慧勇等[28]设计了一种"钻石型"涡流发生器,在FL-31常规高超声速风洞中,利用脉动压力传感器、红外热图技术等对壁面热流进行测量来判断转捩区域,得到了不同涡流发生器高度对转捩区域的影响规律;赵俊波等[29]、张红军等[30]提出了基于T-S波亚谐频共振原理的进气道边界层控制方法,并通过二元进气道实验验证了该方法的有效性,该转捩装置能够大幅减少热流量.由于高超声速飞行的马赫数高,以高空30 km、马赫数为6为例,边界层内涡流发生器的驻点温度理论上可达1771 K,而马赫数为8驻点温度则可达到2981 K,对热防护提出了严重挑战;此外,通常涡流发生器为固定不可调节机构,因此无法根据飞行条件及控制效果反馈控制.
利用外加电磁场对边界层气体作用改变流场状态,相关研究最早可以追溯到20世纪60年代,Rossow[31-32]开展了对导电流体边界层及气流特性的影响研究,但由于当时对转捩的认识不清楚、需求也不明确,且人工电离技术水平以及磁场设备重量大等因素制约,相关研究没有进一步深入开展.
边界层等离子体/磁流体转捩控制技术,具有尺寸小、重量轻、无需移动部件、频带宽、响应迅速等优点,被认为在航空航天领域具有很好的应用前景.其原理是通过在边界层内气体周期性放电产生等离子体,并在磁场的作用下实现对边界层气体的周期性扰动,产生沿流向运动的涡流,促进边界层转捩.
3.2 主要研究进展
对于等离子体/磁流体控制转捩方面的研究,Kimmel[33]在高超声速边界层转捩综述中叙述了Cheng等[34]、Palm等[35]研究了定常磁流体作用对超声速边界层的影响,Cheng等采用DNS分析了磁流体作用可能对第二模态波产生影响,该计算采用Ma=4.5均匀电导率假设;Palm等采用RF射频预电离Ma=4氦气流,测量分析了洛伦兹体积力对壁面静压的影响.2010年,严红等[36-37]将等离子体热激励简化为局部热壁条件($T_{\rm w}$=1.76$T_{\rm ad})$,采用DNS研究了等离子体热激励对Ma=1.5层流边界层转捩的作用机理.研究发现：当采用定常激励时,其对流场的扰动作用在激励区域下游将逐渐耗散;当采用100 kHz非定常激励时,产生的涡扰动将沿流向方向增长,并具有明显的三维特性,形成了发卡涡结构并且转捩过程与K型转捩相似.2014年,Polivanov等[38]也进行了相关研究,他们认为SDBD激励器的体积力的作用较小可以忽略不计,并将等离子体激励简化局部能量源项.采用大涡模拟(LES)研究了激励功率密度、频率等参数对Ma=1.5, 2边界层转捩的影响规律.研究发现,当激励频率过高时,由于局部热量来不及通过对流作用向下游扩散,这将导致其作用效果与定常激励等同,此时无法促使边界层转捩;同年,Keller等[39]也采用局部能量源项的方式,研究了定常激励对Ma=3, 5,7层流边界层的作用效果,同样没有发生转捩,并指出应该增大雷诺数Re$_{kk}$,同时激励要以适当的频率施加.
Purdue大学的Schneider课题组从20世纪90年代就开始针对高超声速转捩问题进行了系统研究.1998年,Ladoon等[40]在Ma=4的静风洞内开展了相关研究,在尖锥上安装放电电极并且同样点辉光放电的方式.他通过热线测试得到了人工扰动波的传播过程. 2009年,Borg[41]在Ma=6静风洞BAM6QT内开展了类似研究,并在X-51前体20%缩比模型安装了辉光放电激励器.放电频率选择为30 kHz和100 kHz两种情况,分别对应第一模态和第二模态扰动波.但是得到的测试结果远不及Ladoon等[40]的理想,并且壁面安装的PCB压力传感器没有测量到第二模态人工扰动波,他根据STABL软件稳定性分析结果进行了简要解释,将其原因归结为激励位置100 kHz的扰动波在中性稳定性曲线的外部,所以扰动波迅速衰弱. Houpt等[42-44]开展了部分转捩控制的可行性研究,并设计了一种浅腔型等离子体激励器,其结构类似于PSJ,结构示意图如图5所示,该激励器的优点在于可以明显降低击穿电压,并且等离子体源的频率达到了100 kHz. 实验在Ma=4.5的风洞中进行,采用光学测试设备波前传感器,可以对边界层内流体的密度脉动进行直接测量,从原理上证明了高频等离子体激励可以对高超声速边界层施加有效扰动.
图5
新窗口打开|下载原图ZIP|生成PPT图5浅腔型等离子体激励结构示意图
Fig.5Scheme of shallow cavity discharge plasma actuator
李益文提出一种电磁式涡流发生方法,其基本原理如图6所示,利用与壁面齐平布置的电极,在边界层内周期性气体放电,并在磁场的作用下定向运动,实现对边界层气流的周期性扰动,诱导沿流向运动的涡流.与传统的涡流发生器相比具有突出优势,由于耐高温电极与壁面齐平,热防护性好;电磁式转捩控制方法可以根据飞行条件,对放电强度和放电频率等参数进行调节,具有激励频带宽、调节方便迅速的特点,并可根据转捩控制效果进行反馈控制.
图6
新窗口打开|下载原图ZIP|生成PPT图6电磁式涡流发生器结构示意图
Fig.6Scheme of electromagnetic vortex generation
采用PIV对静止空气条件下电磁式涡流发生器的流场特性进行了测量,典型条件下的测试结果如图7所示,由于放电过程发光较强,电磁式涡流发生器部分采取了遮蔽处理(图中左下角部位),从图可以看出在磁式涡流发生器的射流速度约1~2 m/s左右,射流在壁面附近诱导形成了漩涡.
图7
新窗口打开|下载原图ZIP|生成PPT图7PIV测试得到的流场特性图
Fig.7Characteristics of flow field by PIV
采用喷管与实验段直连方式,利用小型吸气式超声速风洞(Ma=2),开展了超声速气流条件下的放电特性测试及压力测试研究,实验设备的安装如图8所示,实验段入口截面尺寸30 mm×40 mm,喷管后的试验段采用0.5$^\circ$的边界层修正,电磁式涡流发生器安装在实验段侧面,电极头部间距8 mm,尾部间距12 mm,扩张角为2.9$^\circ$(电极夹角为5.8$^\circ$),涡流发生器前后壁面安装7个动态压力传感器测量边界层动态压力,传感器编号顺气流方向为1~7号.
图8
新窗口打开|下载原图ZIP|生成PPT图8超声速喷管
Fig.8Supersonic nozzle
由于电磁干扰对压电式动态压力传感器造成较大的影响,对压力数据进行处理,边界层动态压力频谱特性如图9所示,红色是电磁式涡流发生器工作,从放电前后曲线对比可以看出,该喷管的边界层脉动主频在约10 K附近,电磁式涡流发生器开启后,功率谱在10 K附近发生较显著的变化,表明边界层湍流度相比得到增强,初步验证了电磁式涡流发生器的可行性,后续研究将利用光学测试手段开展更详细和深入的研究.
图9
新窗口打开|下载原图ZIP|生成PPT图9边界层动态压力频谱特性,有等离子体激励(红色),无等离子体激励(蓝色)
Fig.9Spectral characteristics of boundary layer dynamic pressure, with plasma actuation (red), without plasma actuation (blue)
4 等离子体/磁流体建模
等离子体/磁流体流动控制涉及流场、电场、磁场等多物理场耦合,受制于实验成本与现有测试技术的发展,要深入探究其机理,获得由于实验手段限制而难以获得的细致流场结构,需要采取实验研究与数值模拟并行的方式.4.1 等离子体模型
由于等离子体与流体在时间与空间尺度上的较大差异,等离子体与流体动力学的耦合计算是工程计算中的重难点.目前常用的耦合方法有两种：(1)将等离子体与流体方程在相同的网格和时间尺度上直接耦合计算;(2)假设放电期间气体压力,温度等流场参数不变,求解等离子体动力学方程,并将结果作为流体方程求解的边界条件或源项.等离子体模拟通过求解泊松方程、粒子输运方程、光电离方程和反应方程可以较为准确地描述放电过程.较为常见的有粒子群-蒙特卡罗(PIC-MCC)模型、两方程迁移-扩散模型以及化学反应模型与两方程模型结合的三方程模型.目前针对表面介质阻挡等离子体激励器(SDBD)的数值模拟较多,Soloviev等[45-47]对SDBD进行了较为系统的研究;Shang等[48-49]采用迁移-扩散模型对了高超声速气流下的直流辉光放电过程进行模拟,并通过霍尔参数引入外加磁场的影响,数值模拟结果与实验结果的吻合较好,如图10所示.
图10
新窗口打开|下载原图ZIP|生成PPT图10有无磁场时的放电图像与电子密度对比
Fig.10The comparison of Electron density and discharge images with or without externally applied magnetic fields
虽然等离子体与流体动力学的直接耦合计算可以较为科学地描述等离子体流动控制的物理过程,但其巨大的计算量对于实际工程问题是无法接受的.因此,基于等离子体-流体的单向耦合的研究较多,许多团队使用了开源代码或者商业软件,包括OpenFOAM[50],COMSOL Multi-physics等[51-52],这些模拟工作将等离子体与流体方程置于统一的数据结构框架内以获得对等离子体流动控制自洽的描述.随着对放电物理过程的认识不断深入,有力地推动了唯象学模型的发展.
对于高速流场,等离子体热效应作用机理占主导地位.对于纳秒DBD放电方式,Popov[53]通过化学反应动力分析得到了工程较为关心的快速气体加热效率系数$\eta$$_{\rm total}$,即快速气体加热能量占总放电沉积能量的比率.Takashima等[54]采用唯像学模型,假设流注头部电场是由波前的空间电荷引起的,将其近似为介质表面的一条带点导线,如图11(a)所示.由此,描述二维表面介质阻挡放电的漂移-扩散方程可以被简化为准一维方程系统,在知道电场、电子密度分布和等离子体层厚度随电离波传播的规律时,该方程系统可以获得解析解,计算得到的热源分布规律如图11(b)所示.
图11
新窗口打开|下载原图ZIP|生成PPT图11(a)表面介质阻挡放电唯像模型;(b)维象模型计算所得能量密度分布
Fig.11(a)Schematic of a SDBD phenomenological model; (b)The distribution of energy density computed by SDBD phenomenological model
相比于表面介质阻挡放电,适用于表面辉光,火花或电弧放电的唯象学模型较少,一般将等离子体作用范围将根据实验结果简化为长方体区域,但由于负辉区、正柱区、阳极区等的存在,将导致电导率非均匀分布,所以通常假设电导率符合高斯[55]或超高斯分布[56].
4.2 磁流体模型
磁流体理论模型由流体力学守恒方程组、Maxwell方程组以及磁流体本构方程组成,因此磁流体动力学模型原始形式的数学特性高度复杂.事实上,磁流体动力学是一多尺度问题,例如,空气中扰动的传播速度为声速,而电磁波传播速度可与光速比拟,二者量级的巨大差异所引起的刚性问题在当前尚难逾越,因此并未见到直接对上述原始理论模型进行数值求解的公开文献.在现有计算方法和计算机条件下开展数值模拟研究,必须针对实际问题的特性对理论模型进行适当的简化.磁流体动力学方程组根据磁雷诺数的量级差别可分为八方程模型和五方程模型.因此,针对不同的模型方程,其数值模拟研究所用的方法也不尽相同.对于八方程模型的方程组的求解存在两个数学难题,即刚性问题和边界条件.其中刚性问题源于磁流体通道中磁雷诺数和黏性雷诺数的量级差别过大;边界条件问题则源于出现磁场的散度边界条件,在数学上难于处理.但对于实际飞行条件下的磁流体流动控制问题一般为低磁雷诺数流动.事实上磁雷诺数反映了感应磁场和原有磁场的对比,当磁雷诺数很小时,感应磁场相对于外加磁场可以忽略不计,此时在Navier-Stokes方程组中出现的有关磁场的作用项可以作为彻体力处理.方程组中的变量除了流体力学中的五个变量外,还出现了一个新变量,即电场强度,为了令模型方程组封闭,正确的方法是引进电势函数,导出电势方程.该方程与Navier-Stokes方程组联立构成求解低磁雷诺数下磁流体流动的模型方程[57-58].国内方面,吕浩宇等[59-60]、田正雨等[61]磁流体动力学模型修正与计算方法改进方面做出的积极贡献,推动了国内MHD流动控制的研究发展.
5 关键技术及发展趋势分析
自等离子体/磁流体流动控制这一新概念主动流动控制技术的提出伊始,大量的理论研究成果与原理性实验都证明了其可行,但要实现工程化应用,仍然任重而道远,然而从长远发展来看,由于等离子体/磁流体流动控制技术具有主动性、无运动部件、快速响应、激励频带宽等突出优势,其在超声速特别是高超声速领域具有广阔的应用前景,其未来发展趋势有以下几个方面.5.1 等离子体模型
如何在超声速气流中产生并维持高电导率、均匀稳定的等离子体,一直是制约等离子体/磁流体流动控制技术发展的关键问题,同时提高电离能量利用效率也是关乎其未来能否用于实际飞行器的重要因素.电子束电离是目前理论上效率最高的电离方式,其产生一对电子-离子平均消耗34 eV能量,但由于其需要较为复杂的电子束设备,无法满足机载要求,并且电子束在高密度气流中穿透性较差,目前还无法工程应用;射频放电与微波电离能量利用率较低,一般在几百电子伏特,且电导率较低;高压纳秒脉冲放电是一种非常高效的电离方式,通过瞬间高电场击穿空气提高电离系数,对于空气而言,每个电子平均消耗能量仅为66 eV能量,同时其强电场维持时间远短于放电向不稳定发展时间,因此增强了放电的稳定性,但研究表明,对于空气等电负性气体,由于电子复合较快,导致仅在1μs内电子数密度就衰减了70%以上,而目前的等离子体源技术还难以做到脉冲频率在100 kHz以上,因此很难在时间与空间上产生连续的等离子体.受制于非平衡电离、高频脉冲电源技术的发展限制,目前只采取单一的放电方式无法取得较好的电离效果,需考虑采用脉冲+直流、脉冲+射频等组合放电方式,并且不断提高电源等关键技术,比如减小电源重量、提高脉冲频率等,这些将会对等离子体/磁流体流动控制在未来飞行器上应用提供关键技术支撑.
5.2 改进优化激励参数并探索新型气动激励方式
目前基于等离子体的流动控制气动激励方式有多种形式,包括：有无磁场的表面裸露电极放电、表面介质阻挡放电、等离子体合成射流、大体积磁流体流动控制等.各种激励方式有各自的特点与优势,但同时又有各自的缺点,例如：表面裸露电极放电虽然激励强度较大,但存在能耗较高的缺点;对于表面介质阻挡放电虽然功耗较低,但对于超声速气流其激励强度较弱;等离子体合成射流功耗较大,并且其重频控制效果不佳;大体积磁流体流动控制也受制于现有等离子体源技术的发展,控制效果一般.所以随着等离子体源技术的发展,要不断探索新型的气动激励方式,同时也要在现有技术的基础上,不断优化激励器设计,降低激励功耗,得到最优的电压-电流、脉宽频率等激励参数,并关注其与流场的耦合作用,获得最为合理与高效的控制方式.
6 结 论
本文对进气道等离子体/磁流体流动控制研究进行了综述和分析,主要结论如下：(1)等离子体/磁流体流动控制可能更好的解决在不同马赫数飞行条件下进气道激波系控制中存在的挑战与困难;进气道采用较小马赫数设计,在高马赫数条件下,等离子体/磁流体流动控制可实现激波系外推唇口,该技术有望在需要短时间流动控制的高马赫数导弹上走向工程应用.
(2)边界层等离子体/磁流体转捩控制技术,具有尺寸小、重量轻、无需移动部件、频带宽、响应迅速,可根据飞行条件对控制参数进行调节,并根据转捩效果进行反馈控制的优势和特点,在进气道转捩控制中具有很好的应用前景.
(3)等离子体/磁流体流动控制技术属于多学科交叉,后续发展中需要进一步创新等离子体产生技术及激励方式,建立等离子体与流体的全耦合计算模型,深入探究流动控制过程中的机理,为基础研究及工程应用提供重要的基础.
参考文献 原文顺序
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被引期刊影响因子
DOIURL [本文引用: 1]
URL [本文引用: 1]
介绍了磁流体动力学在航空工程中的主要应用方式,主要包括:磁流体冲压组合发动机、磁流体涡轮组合发动机、燃烧室后磁流体发电、表面磁流体发电、磁流体加速风洞、磁流体推力矢量、进气道大尺寸磁流体流动控制、边界层分离流动控制、边界层转捩控制、飞行器头部热流控制等;探讨了磁流体技术在应用中存在的关键科学与技术问题,对导电流体的产生、磁流体实验设备与实验技术、多场耦合机理及数值模拟方法等进行了分析;最后对磁流体技术在航空工程上的应用与发展进行了总结与展望.
URL [本文引用: 1]
介绍了磁流体动力学在航空工程中的主要应用方式,主要包括:磁流体冲压组合发动机、磁流体涡轮组合发动机、燃烧室后磁流体发电、表面磁流体发电、磁流体加速风洞、磁流体推力矢量、进气道大尺寸磁流体流动控制、边界层分离流动控制、边界层转捩控制、飞行器头部热流控制等;探讨了磁流体技术在应用中存在的关键科学与技术问题,对导电流体的产生、磁流体实验设备与实验技术、多场耦合机理及数值模拟方法等进行了分析;最后对磁流体技术在航空工程上的应用与发展进行了总结与展望.
[本文引用: 1]
Changsha: National University of Defense Technology Press,
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URLMagsci [本文引用: 1]
<p><正> 航空工业从开始以来,一个很重要的目标,就是不断地提高飞行器的速度。因为速度的增加,气流里所产生的现象,就逐渐复杂起来。二十几年前,飞行速度平均在每小时三四百公里左右,建立在空气是不可压缩没有粘性的假设上的流体力学,对于飞机的设计,就有了很大的贡献。后来为了战争的要求,飞行的速度提到六七百公里,在飞行和制造上,就第一次发生了困难,就是所谓''空气可压缩的困难"。</p>
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Abstract. This work continues a series of [1,2] initiated under AJAX program [3] with a purpose to study the external effects on the shock-wave structures arising in a diffuser with a total inner flow compression. The main part of the diffuser is a linearly convergent channel. The shock-wave structure incorporates two attached shocks affected by each other in the diffuser. External electric field arranged in such a manner that it enhanced a magneto-induced current. A flow is a result of the MHD interaction itself as well as a gas heating under the external electric field.
DOIURL
react-text: 277 In the work pulsed interaction between an ionized gas flow and magnetic and electric fields was studied. The experiments and numerical simulation were carried out in a diffuser with total internal flux compression with a Mach number of = 4.3 at the diffuser inlet. The interaction pulse was determined by the pulse of electric current between electrodes located in the inlet part of the diffuser,... /react-text react-text: 278 /react-text [Show full abstract]
DOIURL [本文引用: 1]
ABSTRACT The paper examines the possibility of controlling scramjet inlets in off-design conditions by operating a near-surface MHD system upstream of the inlet. The required electrical conductivity in air is supposed to be created by electron beams injected into the air from the vehicle along magnetic field lines. A simple model of beam-generated ionization profile is developed and coupled with plasma kinetics, MHD equations, and 2D inviscid flow equations. Calculations show that an MHD system with reasonable parameters could bring shocks back to the cowl lip when flying at Mach numbers higher than those for which the inlet was optimized. The MHD effect is not reduced to heating only, as the work by j B forces is a substantial part of the overall effect. Power requirements for ionizing e-beams could be lower than the electrical power extracted with MHD, so that a net power would be generated on board. Problems associated with high Hall fields are discussed.
DOIURL [本文引用: 1]
The paper is devoted to analysis of magnetohydrodynamic (MHD) control of forebody flow compression, shock incidence, and air mass capture increase in scramjet inlets for vehicles that would fly at Mach 5-10. Due to the low static temperature, nonequilibrium electrical conductivity is created by electron beams injected into the gas along magnetic field lines. Two-dimensional inviscid steady-state flow equations are solved jointly with equations describing electron beam-induced ionization profiles, plasma kinetics, and MHD equations. Several scenarios are considered, based in part on earlier work by the present authors. The scenario with an on-ramp MHD accelerator that should increase mass capture at Mach numbers lower than the design value has only disadvantages, since the MHD device consumes high power, reduces total pressure, and actually decreases mass capture due to Joule heating and thermal expansion of the gas. A modest increase in mass capture can in principle be accomplished in a Faraday MHD generator mode, if the magnetic field has components both parallel and orthogonal to the flow. However, this scenario involves very large volumes of strong magnetic fields, and the mass capture increase is due mostly to a nonuniform gas heating. In a new "virtual cowl" scenario, a localized off-body energy addition is used to increase mass capture at Mach numbers lower than the design value, while also even increasing total pressure at the inlet. The principal focus of this paper is on inlet control at flight Mach higher than the design value. The shocks that would otherwise enter the inlet can be moved back to the cowl lip by placing an MHD generator at one of the compression ramps. Calculations and qualitative arguments show that the best performance of such a device (minimal losses of total pressure) is achieved with a very short MHD region in conjunction with high-current ionizing electron beam; the MHD region should be placed as far upstream (close to the vehicle nose) as possible. An MHD energy bypass scenario with on-ramp MHD generator for inlet control is briefly discussed.
DOIURL [本文引用: 1]
为了开展磁流体(MHD)流动控制原理研究,建立了磁流体技术试验系统,采用电容耦合射频-直流组合放电对Ma=3.5气流进行电离,在磁场作用下产生顺/逆气流方向的洛伦兹力控制流场,采用试验段静压变化来监测磁流体流动控制效果,通过一维模型计算磁流体流动控制过程中流场变化情况,分析磁流体流动控制效果;通过添加电磁源项的Navier-Stokes方程耦合电势泊松方程建立了二维磁流体动力模型,对磁流体流动控制进行数值模拟研究.主要结论如下:在磁场约束下,电容耦合射频-直流组合放电能够在Ma=3.5流场中产生大体积均匀电流,电导率约0.015 S/m;在焦耳热和洛伦兹力作用下,磁流体加速时静压升高了130 Pa,减速时静压升高了200 Pa;磁流体流动控制过程中,仅有不足10％的能量在磁流体通道内发生了作用;数值模拟结果显示,在试验条件下,加速时静压升高了128 Pa,减速时静压升高了208 Pa,与试验结果基本吻合.
DOIURL [本文引用: 1]
为了开展磁流体(MHD)流动控制原理研究,建立了磁流体技术试验系统,采用电容耦合射频-直流组合放电对Ma=3.5气流进行电离,在磁场作用下产生顺/逆气流方向的洛伦兹力控制流场,采用试验段静压变化来监测磁流体流动控制效果,通过一维模型计算磁流体流动控制过程中流场变化情况,分析磁流体流动控制效果;通过添加电磁源项的Navier-Stokes方程耦合电势泊松方程建立了二维磁流体动力模型,对磁流体流动控制进行数值模拟研究.主要结论如下:在磁场约束下,电容耦合射频-直流组合放电能够在Ma=3.5流场中产生大体积均匀电流,电导率约0.015 S/m;在焦耳热和洛伦兹力作用下,磁流体加速时静压升高了130 Pa,减速时静压升高了200 Pa;磁流体流动控制过程中,仅有不足10％的能量在磁流体通道内发生了作用;数值模拟结果显示,在试验条件下,加速时静压升高了128 Pa,减速时静压升高了208 Pa,与试验结果基本吻合.
DOIURL
At the present time numbers of theoretical and experimental works have been published, which show that energy release to airflow near/fore streamlined bodies can reduce a total drag of these bodies. It occurs at high level of energetic efficiency (sometimes, much more than 1) [1-5]. Several works are described efforts in a field of plasma influence on a viscous friction and separation zone... [Show full abstract]
DOIURL
Abstract The results of experimental and numerical investigations of the interaction between the near-wall electrical discharge and supersonic airflow in an aerodynamic channel with constant and variable cross sections are presented. Peculiar properties of the surface quasi-direct-current discharge generation in the flow are described. The mode with flow separation developing outside the discharge region is revealed as a specific feature of such a configuration. An interaction model is proposed on the basis of measurements and observations. A regime of gas-dynamic screening of a mechanical obstacle installed on the channel wall is studied. Variation of the main flow parameters caused by the surface discharge is quantified. The ability of the discharge to shift an oblique shock in an inlet is demonstrated experimentally. The influence of relaxation processes in nonequilibrium excited gas on the flow structure is analyzed. Comparison of the experimental data with the results of calculations based on the analytical model and on numerical simulations is presented.
DOIURL
The paper is aimed to present the results of experimental and computational research in a field of supersonic Flow Control by means of near-surface electrical discharge generation. The specific task of this activity is to demonstrate the steering effect of low-temperature nonequilibrium plasma on supersonic flow structure in a compression ramp. The experiments were arranged in connected pipe configuration (lab-scale) at M=2-2.5. CFD efforts in 3D NS approach clarify extra features of plasma-flow-surface interaction.
DOIURL
Abstract The paper presents results of experimental and computational study of near-surface electrical discharge effect on shocks configuration in a compression ramp and 2D inlet performance in M=2-2.5 supersonic flow. Two models were designed for experiments in connecting pipe configuration and in free-stream correspondingly. The models contain a special insertion with flush mounted plasma generator arranged ahead of the first rupture of two-corner ramp. It is shown that the plasma deposition allows an accurate regulation of shock angle and reflection line position on opposite wall of the first model and in vicinity of cowl lip of the second model. The magnitude of regulation effect depends on the power release. The CFD efforts in 3D NS approach clarify some extra features of plasma-flow-surface interaction.
DOIURL [本文引用: 1]
The objective of this work was to study the steering effect of a weakly ionized plasma on a supersonic flow structure in a two-dimensional aerodynamic configuration with a three-shock compression ramp in an off-design operational mode. Experiments were performed in wind tunnel T-313 of ITAM SB RAS, with the model air inlet designed for operation at a flow of Mach number M 02=022. The inlet was tested at M 02=022, 2.5, and 3 and with Re 02=02(25–36)02×0210 6 /m and an angle of attack AoA02=020°, 5°, and 8°. For the regulation of the inlet characteristics, a plasma generator with electrical power W pl 02=022–1002kW was flush-mounted upstream of the compression ramp. A significant plasma effect on the shock configuration at the inlet and on the flow parameters after air compression is considered. It is shown that the main shock wave angle is controllable by means of the plasma power magnitude and, therefore, can be accurately adjusted to the cowl lip of an inlet with a fixed geometry. An additional plasma effect has been demonstrated through a notable increase in the pressure recovery coefficient in a flowpass extension behind the inlet because of an nearly isentropic pattern of flow compression with the plasma turned on. Numerical simulation brings out the details of 3D distribution of the flow structure and parameters throughout the model at thermal energy deposition in inlet near the compression surfaces. We conclude that the plasma-based technique may be a feasible method for expanding supersonic inlet operational limits.
DOIURL [本文引用: 1]
ABSTRACT This paper reviews work underway on the development of new technologies that will enhance the performance of hypersonic vehicles through plasma-related processes and the utilization of MHD to provide the large power levels that are required to drive these processes. Three technologies are discussed for plasma-enhanced hypersonic performance. These include: (1) the use of off-body plasmas for drag reduction, steering, and enhanced inlet performance; (2) the use of surface or near-surface plasmas for mitigating local heating and controlling separation; and (3) the use of electron beam and plasma processes for controlling combustion for enhanced performance inside the engine and for local heat addition applications in other regions of the flow path. For realistic scale vehicles, the energies required for these applications exceed the present capability of on-board auxiliary power units, and, therefore, will require power to be generated directly from the hypersonic air passing over the vehicle or through the engine. In the high Mach number regime characteristic of re-entry vehicles, there is sufficient heating of the air to allow MHD power extraction using equilibrium ionization of alkali vapor seed material. By replacing a portion of the vehicle surface with a hollow core truss structure containing an embedded magnet coil, hundreds of kilowatts of power can be extracted during re-entry and used for vehicle control or other applications. At lower Mach numbers, MHD power extraction can be done downstream of the engine, then the temperature of the exhaust can be high enough to allow conductivity to be achieved with alkali seeding. This MHD generated power extracted from the flow aft of the engine can be used for plasma control upstream of the engine as well as for engine performance enhancement. Some aspects of this reverse energy bypass concept are analyzed in the paper, including plasma heating of the inlet flow that would allow elimination of the isolator, snowplow surface arcs for boundary layer and separation control, and electron beam and microwaves for initiation and control of combustion.
DOIURL [本文引用: 1]
This project was a theoretical and experimental research effort on the use of MHD body forces and plasmas for boundary layer control and power extraction in supersonic flow, and on the development of new diagnostics for plasmas and for high-speed flows. The first part of this final report addresses MHD processes for control and power extraction. This section includes the constricted DC driven... [Show full abstract]
DOIURL [本文引用: 1]
This study demonstrates the potential for shockwave–turbulent boundary layer interaction control in air using low current DC constricted surface discharges forced by moderate strength magnetic fields. An analytical model describing the physics of magnetic field forced discharge interaction with boundary layer flow is developed and compared to experiments. Experiments are conducted in a Mach 2.6 indraft air tunnel with discharge currents up to 30002mA and magnetic field strengths up to 502 Tesla . Separation- and non-separation-inducing shocks are generated with diamond-shaped shockwave generators located on the wall opposite to the surface electrodes, and flow properties are measured with schlieren imaging, static wall pressure probes and acetone flow visualization. The effect of plasma control on boundary layer separation depends on the direction of the Lorentz force ( j × B ). It is observed that by using a Lorentz force that pushes the discharge upstream, separation can be induced or further strengthened even with discharge currents as low as 3002mA in a 3- Tesla magnetic field. If shock-induced separation is present, it is observed that by using Lorentz force that pushes the discharge downstream, separation can be suppressed, but this required higher currents, greater than 8002mA. Acetone planar laser scattering is used to image the flow structure in the test section and the reduction in the size of recirculation bubble and its elimination are observed experimentally as a function of actuation current and magnetic field strength.
DOIURL [本文引用: 1]
A magnetically driven DC surface plasma discharge has shown promise for supersonic boundary layer control. Previous experiments in supersonic flow performed using this plasma actuator have shown significant control over the shockwave- turbulent boundary layer interaction region. In this study, the velocity of the surface jet in quiescent air is measured and compared with predictions. A time gated schlieren technique is used to visualize the jet and determine the velocity. The surface plasma column appears as a transverse "arc" between two slightly diverging electrodes and is driven by j x B forces so that it sweeps the neutral gas near the surface creating a surface jet.The jet velocity generated by the magneto-gas dynamic actuator is measured in a 1 Tesla magnetic field at different actuation currents below 100mA in air at pressures in the range of 100 - 400 Torr
DOIURL [本文引用: 1]
ABSTRACT The physics of constricted moving nonequilibrium discharges between surface-imbedded electrodes in strong magnetic field is analyzed. In the experiments reported elsewhere, the almost-parallel electrodes were slightly diverging in either upstream or downstream direction, and the discharge was confined to the boundary layer of supersonic flow. The magnetic field perpendicular to the surface caused the discharge to move downstream or upstream with high velocity. Electron and ion number densities and the discharge radius are estimated. Based on the equations of motion for electrons and ions, the discharge propagation velocity is calculated and is found to be in agreement with experimental data. Boundary layer velocity increase caused by the moving discharge is calculated, and ways to enhance the gas acceleration and the ratio of "push power" to the heating rate are analyzed. Possible mechanisms of discharge canting and its spiral structure are discussed.
DOIURL [本文引用: 1]
The Navier Stokes equations were solved using a high-fidelity time-implicit numerical scheme and an implicit large-eddy simulation approach to investigate plasma-based flow control for supersonic flow over a compression ramp. The configuration included a flat-plate region to develop an equilibrium turbulent boundary layer at Mach 2.25, which was validated against a set of experimental measurements. The fully turbulent boundary-layer flow traveled over a 24 deg ramp and produced an unsteady shock-induced separation. A control strategy to suppress the separation through a magnetically-driven surface-discharge actuator was explored. The size, strength, and placement of the model actuator were based on recent experiments at the Princeton University Applied Physics Group. Three control scenarios were examined: steady control, pulsing with a 50% duty cycle, and a case with significant Joule heating. The control mechanism was very effective at reducing the time-mean separation length for all three cases. The steady control case was the most effective, with a reduction in the separation length of more than 75%. The controller was also found to significantly reduce the low-frequency content of the turbulent kinetic energy spectra within the separated region and reduce the total turbulent kinetic energy downstream of reattachment.
DOIURLMagsci [本文引用: 1]
对MHD(mechanisms of magnetohy drodynamics)控制超声速平板湍流边界层的机理进行了理论研究和数值模拟. 理论上,采用等离子体低频近似碰撞频率模型，建立等离子体中电子和离子的力平衡方程，得到等离子体速度、极化电场以及边界层速度. 数值上，通过空间HLLE格式、LU--SGS时间推进求解时均磁流体动力学湍流方程，其中湍流模型采用sst--k\omega双方程模型. 研究结果表明：(1)边界层速度的理论结果和数值结果误差在7%范围内；(2)只有磁场而电场为零时，洛仑兹力起到减小摩阻的作用. 施加电场后，洛仑兹力能够加速边界层低速区流体；(3) 在边界层外层，越靠近壁面，作用参数越小;而在边界层近壁区黏性底层，虽然惯性力减小， 但黏性力却迅速增加，因此越靠近壁面，作用参数反而越大，加速低速流的代价增加.
DOIURLMagsci [本文引用: 1]
对MHD(mechanisms of magnetohy drodynamics)控制超声速平板湍流边界层的机理进行了理论研究和数值模拟. 理论上,采用等离子体低频近似碰撞频率模型，建立等离子体中电子和离子的力平衡方程，得到等离子体速度、极化电场以及边界层速度. 数值上，通过空间HLLE格式、LU--SGS时间推进求解时均磁流体动力学湍流方程，其中湍流模型采用sst--k\omega双方程模型. 研究结果表明：(1)边界层速度的理论结果和数值结果误差在7%范围内；(2)只有磁场而电场为零时，洛仑兹力起到减小摩阻的作用. 施加电场后，洛仑兹力能够加速边界层低速区流体；(3) 在边界层外层，越靠近壁面，作用参数越小;而在边界层近壁区黏性底层，虽然惯性力减小， 但黏性力却迅速增加，因此越靠近壁面，作用参数反而越大，加速低速流的代价增加.
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高超声速附面层的转捩预测一直是流体力学研究中的难点，转捩前后物面的摩擦系数和传热系数会发生改变，转捩位置的不同会影响到飞行器表面热环境，进而使得飞行器的气动弹性特性发生显著变化.鉴于高超声速附面层转捩预测的不确定性，本文分析了转捩位置对高超声速全动舵面热气动弹性的影响.首先分别用层流模型和湍流模型求解N-S方程,得到气动热环境，并对气动热进行参数化；然后在不同转捩位置情况下构造出不同转捩位置的热分布模型，基于此种温度分布，结合热应力和材料属性的影响分析结构的热模态，将结构模态插值到气动网格上，采用基于CFD的当地流活塞理论进行气动弹性分析.以M=6，H=15km的某舵面为对象进行计算,结果表明：(1)随着转捩位置向后缘移动，结构频率上升，结构颤振速度呈增大趋势，转捩位置的变化能够带来颤振临界速度最大6%的变化量；(2)当转捩位置位于舵轴附近时，结构的颤振特性变化剧烈.通过刚度特性的分解和分析发现,导致颤振特性变化的主要因素在于舵轴的刚度特性变化，舵轴的影响量占整个结构刚度特性变化量的80%以上.
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高超声速附面层的转捩预测一直是流体力学研究中的难点，转捩前后物面的摩擦系数和传热系数会发生改变，转捩位置的不同会影响到飞行器表面热环境，进而使得飞行器的气动弹性特性发生显著变化.鉴于高超声速附面层转捩预测的不确定性，本文分析了转捩位置对高超声速全动舵面热气动弹性的影响.首先分别用层流模型和湍流模型求解N-S方程,得到气动热环境，并对气动热进行参数化；然后在不同转捩位置情况下构造出不同转捩位置的热分布模型，基于此种温度分布，结合热应力和材料属性的影响分析结构的热模态，将结构模态插值到气动网格上，采用基于CFD的当地流活塞理论进行气动弹性分析.以M=6，H=15km的某舵面为对象进行计算,结果表明：(1)随着转捩位置向后缘移动，结构频率上升，结构颤振速度呈增大趋势，转捩位置的变化能够带来颤振临界速度最大6%的变化量；(2)当转捩位置位于舵轴附近时，结构的颤振特性变化剧烈.通过刚度特性的分解和分析发现,导致颤振特性变化的主要因素在于舵轴的刚度特性变化，舵轴的影响量占整个结构刚度特性变化量的80%以上.
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在30～70 km空域机动飞行的高超声速飞行器的优点是可以耦合利用所处空域的空气产生的升力和高速飞行的离心力进行远距离机动滑翔飞行,具有重要的实用价值.尽管过去数十年在高超声速流动研究方面取得显著进展,但在设计研究近空间远程滑翔的高超声速飞行器方面仍然存在许多挑战,特别是对特定飞行条件下的流动机理了解不清楚.本文介绍了作者研究团队在开展近空间高超声速飞行器有关的关键气动问题方面的研究进展,主要包括:建立了近空间高超声速飞行的流动模型,发展了系统的相关计算空气动力学方法,针对高空高速飞行条件下稀薄气体效应和真实气体效应的耦合作用影响研究了合适的滑移边界条件,考虑了不同组分存在条件下的温度、速度和压力的滑移效应影响;提出了飞行器气动外形的动态优化方法,获得了可工程实用化的高升阻比飞行器气动外形;建立了高速飞行器动稳定性理论,在实现高超声速飞行器动态稳定飞行方面取得重大进展;最后讨论了高超声速飞行器设计中进一步需要关注的若干关键技术和科学问题、可能解决的途径及其所涉及的学科发展方向.
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在30～70 km空域机动飞行的高超声速飞行器的优点是可以耦合利用所处空域的空气产生的升力和高速飞行的离心力进行远距离机动滑翔飞行,具有重要的实用价值.尽管过去数十年在高超声速流动研究方面取得显著进展,但在设计研究近空间远程滑翔的高超声速飞行器方面仍然存在许多挑战,特别是对特定飞行条件下的流动机理了解不清楚.本文介绍了作者研究团队在开展近空间高超声速飞行器有关的关键气动问题方面的研究进展,主要包括:建立了近空间高超声速飞行的流动模型,发展了系统的相关计算空气动力学方法,针对高空高速飞行条件下稀薄气体效应和真实气体效应的耦合作用影响研究了合适的滑移边界条件,考虑了不同组分存在条件下的温度、速度和压力的滑移效应影响;提出了飞行器气动外形的动态优化方法,获得了可工程实用化的高升阻比飞行器气动外形;建立了高速飞行器动稳定性理论,在实现高超声速飞行器动态稳定飞行方面取得重大进展;最后讨论了高超声速飞行器设计中进一步需要关注的若干关键技术和科学问题、可能解决的途径及其所涉及的学科发展方向.
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高超声速激波与湍流边界层干扰会导致飞行器表面出现局部热流峰值,严重影响飞行器气动性能和飞行安全.针对高马赫数激波干扰问题,以往数值研究多采用雷诺平均方法,而在直接数值模拟方面的相关工作较为少见.开展高超声速激波与湍流边界层干扰的直接数值模拟研究,有助于进一步提升对其复杂流动机理认识和理解,同时也将为现有湍流模型和亚格子应力模型的改进提供理论依据.采用直接数值模拟方法对来流马赫数6.0,34?压缩拐角内激波与湍流边界层的干扰问题进行了研究.基于雷诺应力各向异性张量,分析了高超声速湍流边界层在压缩拐角内的演化特性.通过对湍动能输运方程的逐项分析,系统地研究了可压缩效应对湍动能及其输运的影响机制.采用动态模态分解方法,探讨了干扰流场的非定常运动历程.研究结果表明,随着湍流边界层往下游发展,近壁湍流的雷诺应力状态由两组元轴对称状态逐渐演化为两组元状态,外层区域则由轴对称膨胀趋近于各向同性.干扰流场内存在强内在压缩性效应(声效应),其对湍动能输运的影响主要体现在压力-膨胀项,而对膨胀-耗散项影响较小.高超声速下压缩拐角内的非定常运动仍存在以分离泡膨胀/收缩为特征的低频振荡特性,其物理机制与分离泡剪切层密切相关.
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高超声速激波与湍流边界层干扰会导致飞行器表面出现局部热流峰值,严重影响飞行器气动性能和飞行安全.针对高马赫数激波干扰问题,以往数值研究多采用雷诺平均方法,而在直接数值模拟方面的相关工作较为少见.开展高超声速激波与湍流边界层干扰的直接数值模拟研究,有助于进一步提升对其复杂流动机理认识和理解,同时也将为现有湍流模型和亚格子应力模型的改进提供理论依据.采用直接数值模拟方法对来流马赫数6.0,34?压缩拐角内激波与湍流边界层的干扰问题进行了研究.基于雷诺应力各向异性张量,分析了高超声速湍流边界层在压缩拐角内的演化特性.通过对湍动能输运方程的逐项分析,系统地研究了可压缩效应对湍动能及其输运的影响机制.采用动态模态分解方法,探讨了干扰流场的非定常运动历程.研究结果表明,随着湍流边界层往下游发展,近壁湍流的雷诺应力状态由两组元轴对称状态逐渐演化为两组元状态,外层区域则由轴对称膨胀趋近于各向同性.干扰流场内存在强内在压缩性效应(声效应),其对湍动能输运的影响主要体现在压力-膨胀项,而对膨胀-耗散项影响较小.高超声速下压缩拐角内的非定常运动仍存在以分离泡膨胀/收缩为特征的低频振荡特性,其物理机制与分离泡剪切层密切相关.
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Active and passive methods for tripping hypersonic boundary layers have been examined in NASA Langley Research Center wind tunnels using a Hyper-X model. This investigation assessed several concepts for forcing transition, including passive discrete roughness elements and active mass addition (or blowing), in the 20-Inch Mach 6 Air and the 31-Inch Mach 10 Air Tunnels. Heat transfer distributions obtained via phosphor thermography, shock system details, and surface streamline patterns were measured on a 0.333-scale model of the Hyper-X forebody. The comparisons between the active and passive methods for boundary layer control were conducted at test conditions that nearly match the Hyper-X nominal Mach 7 flight test-point of an angle-of-attack of 2-deg and length Reynolds number of 5.6 million. For passive roughness, the primary parametric variation was a range of trip heights within the calculated boundary layer thickness for several trip concepts. The passive roughness study resulted in a swept ramp configuration, scaled to be roughly 0.6 of the calculated boundary layer thickness, being selected for the Mach 7 flight vehicle. For the active blowing study, the manifold pressure was systematically varied (while monitoring the mass flow) for each configuration to determine the jet penetration height, with schlieren, and
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在FL-31高超声速风洞分别开展了进气道的自然转捩和强制转捩风洞试验,试验Ma数为5、6和7,迎角为1°。通过红外热图得到了壁面的热流分布,从中得到了转捩区域。强制转捩装置为钻石型涡流发生器。随着涡流发生器高度的增加,强制转捩区域逐渐前移,得到了涡流发生器的有效高度,实现了强制转捩的目的。
DOIURL [本文引用: 1]
在FL-31高超声速风洞分别开展了进气道的自然转捩和强制转捩风洞试验,试验Ma数为5、6和7,迎角为1°。通过红外热图得到了壁面的热流分布,从中得到了转捩区域。强制转捩装置为钻石型涡流发生器。随着涡流发生器高度的增加,强制转捩区域逐渐前移,得到了涡流发生器的有效高度,实现了强制转捩的目的。
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针对超燃冲压发动机进气道由于激波-边界层干扰引起的边界层分离以及进气道堵塞问题,提出了一种基于Tollmien-Schlichting(T-S)波谐频共振原理的进气道边界层控制方法,并通过一种典型的二元进气道风洞试验进行了方法验证.研究结果表明,提出的转捩控制方法能够较好地消除由激波-边界层干扰而引起的边界层分离现象,进而降低边界层分离对进气道性能的不良影响,确保进气道性能.同时说明,在超燃进气道设计过程中,需要充分考虑边界层转捩问题,引入有效的边界层转捩控制方法,以保证发动机的正常工作.
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针对超燃冲压发动机进气道由于激波-边界层干扰引起的边界层分离以及进气道堵塞问题,提出了一种基于Tollmien-Schlichting(T-S)波谐频共振原理的进气道边界层控制方法,并通过一种典型的二元进气道风洞试验进行了方法验证.研究结果表明,提出的转捩控制方法能够较好地消除由激波-边界层干扰而引起的边界层分离现象,进而降低边界层分离对进气道性能的不良影响,确保进气道性能.同时说明,在超燃进气道设计过程中,需要充分考虑边界层转捩问题,引入有效的边界层转捩控制方法,以保证发动机的正常工作.
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为确保高超声速进气道的安全工作,其压缩面边界层在进入其内流道前必须完成转捩。针对高超声速进气道边界层转捩需要,依据二维高超声速边界层转捩机理,尝试了一种新型低阻高效的边界层人工转捩方法,在FD-07风洞中开展了试验验证。试验中首先通过进气道对称面压力分布和激波纹影获得进气道的自起动情况,进而推断进气道入口前的边界层转捩情况。试验包括进气道前体边界层自然转捩和人工转捩,试验结果表明在Ma=5、6,迎角α=0°来流条件下,使用同一波长的人工转捩带可以成功实现进气道边界层转捩,验证了基于线性稳定性理论设计的人工转捩带在宽马赫数范围的适用性。
DOIURL [本文引用: 1]
为确保高超声速进气道的安全工作,其压缩面边界层在进入其内流道前必须完成转捩。针对高超声速进气道边界层转捩需要,依据二维高超声速边界层转捩机理,尝试了一种新型低阻高效的边界层人工转捩方法,在FD-07风洞中开展了试验验证。试验中首先通过进气道对称面压力分布和激波纹影获得进气道的自起动情况,进而推断进气道入口前的边界层转捩情况。试验包括进气道前体边界层自然转捩和人工转捩,试验结果表明在Ma=5、6,迎角α=0°来流条件下,使用同一波长的人工转捩带可以成功实现进气道边界层转捩,验证了基于线性稳定性理论设计的人工转捩带在宽马赫数范围的适用性。
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The use of a magnetic field to control the motion of electrically conducting fluids is studied. The incompressible boundary-layer solutions are found for flow over a flat plate when the magnetic field is fixed relative to the plate or to the fluid. The equations are integrated numerically for the effect of the transverse magnetic field on the velocity and temperature profiles, and hence, the skin friction and rate of heat transfer. It is concluded that the skin friction and the heat-transfer rate are reduced when the transverse magnetic field is fixed relative to the plate and increased when fixed relative to the fluid. The total drag is increased in all of the areas.
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ABSTRACT Neutral stability curves pertaining to a two-dimensional infinitesimal sinusoidal disturbance are presented for the laminar flow of an incompressible, electrically conducting fluid over a semi-infinite flat plate in the presence of either a coplanar or transverse magnetic field. The magnetic field is found to be stabilizing in all of the cases studied except one.
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Las Organizaciones al encontrarse en un mundo altamente competitivo y de turbulencia, están optando en mirarse hacia adentro, en una actitud introspectiva, obviamente, sin descuidar su atención al cliente, o lo que se conoce como la Función Relacional con el Cliente o el Customer Relationshing Managemente (CRM). Por eso es de vital importancia, conocer las Estructura Interna de los Costos de las organizaciones, y el manejo de las mismas, los cuales evidentemente repercutirán en el precio y en la rentabilidad de las organizaciones. Es así que las Empresas están optando por herramientas de planificación en el manejo de los costos, como es el tratamiento de los costos directos o variables, herramienta valiosa que incluye en la estructura de los costos, solo aquellos costos o erogaciones variables en la fabricación del producto o prestación de servicios. Los costos por absorción, que incluyen en la estructura de los costos tanto los costos variables como los fijos, en la determinación de los costos unitarios de producción.
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ABSTRACT This paper investigates, by direct numerical simulation, the effects of an imposed magnetic field on a weakly ionized Mach 4.5 boundary layer. The main emphasis of the study is on MHD effects on the second mode instability in supersonic boundary layer. The imposed magnetic fields are generated by placing twodimensional magnetic dipoles below of the flat plate surface. The gas is assumed to have a constant electrical conductivity of l00mho/m. The magnetic Reynolds number of the flow is small so that the induced magnetic field in the flow is neglected. The governing equations of the MHD flow, which are the Navier-Stokes equations with the applied magnetic force terms, are computed by a fifth-order shock-fitting numerical scheme. A series of cases with different imposed magnetic fields have been investigated on the influences of imposed magnetic field on both the mean flow and on the second mode stability. It is found that the imposed magnetic fields significantly retard the streamwise velocity and reduce the local skin friction in the mean flow. For the case of a strong imposed magnetic field, a local separation region is generated in the mean flow with a strong adverse pressure gradient. Meanwhile, the second mode wave disturbances are found to be stabilized by the imposed magnetic fields, even for the case with strong adverse pressure gradient and a local separated flow region. This unexpected strong stabilization of the second mode wave is believed to be caused by the alteration of the steady base flow by the magnetic field. It is also found that, unlike the second mode, the magnetic fields slightly destabilize the first mode waves because the different instability mechanisms of the two modes. The results presented in this paper are the first concrete results on the interaction of second instability mode with magnetic field in a supersonic boundary layer. 2002 by American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
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This paper investigates the mechanism of steady and unsteady thermal perturbation (also denoted as thermal bump) in a Mach 1.5 flat plate boundary layer. A high-fidelity upwind-biased third-order Roe scheme is used with the compressive van Leer harmonic limiter on a suitably refined mesh. The study consists of two parts. In the first part, the effects of the steady and pulsed thermal bumps are explored. It is shown that the finite-span thermal bumps generate streamwise vortices. With steady heating, the disturbance decays downstream. However, when the thermal bump is pulsed, vortex shedding is observed and the streamwise vortical disturbance grows with downstream distance, consistent with linear stability analysis. The integrated disturbance energy indicates that streamwise kinetic disturbance energy growth dominates over those associated with other two velocity and thermodynamic components. The second part of this paper explores the physical consequences of the nonlinear dynamics between the vortices produced by the pulsed bump and the compressible boundary layer. The resulting three-dimensional flow distortion generates hairpin structures which are aligned in the streamwise direction, suggesting that the transition process bears some similarity to K-type breakdown. The arrangement of these vortices is connected to the low-speed streaks observed in the evolving boundary layer. The shape factor, velocity, and Reynolds stress profiles suggest that the perturbed flow shows initiation of transition to turbulence, but remains transitional at the end of the plate. (C) 2010 American Institute of Physics. [doi:10.1063/1.3432513]
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A three-dimensional numerical study is performed to explore the effect of pulsed spanwise-periodic surface thermal perturbation (also denoted as thermal bump) in a Mach 1.5 flat plate laminar boundary layer. A high-resolution upwind-biased Roe method is used with the compressive Van Leer harmonic limiter on a suitably refined mesh. The dependence of flow stability characteristics on the variation of thermal bump geometry (shape and dimension) and pulsing properties (disturbance amplitude and frequency) is assessed. It is shown that the finite-span thermal bumps generate streamwise vortices. When the thermal bump is pulsed, vortex shedding is observed, and the streamwise vorticity grows with the downstream distance. Analysis of the integrated disturbance energy indicates that the streamwise kinetic disturbance energy dominates over those associated with other two velocity and thermodynamic components. Immediately downstream of the bump, the dominant frequency corresponds to that of the imposed excitation while higher harmonic components are observed farther downstream. An analysis of parametric variation of bump shape and dimension indicates that finite bump span is important in injecting three dimensionality and that the rectangular shape results in faster disturbance growth than the circular one. The study also concludes that disturbance growth is non-linear with bump temperature and has a strong connection with pulsing frequency.
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Boundary-layer flow control is a prerequisite for a safe and efficient operation of future hypersonic transport systems. Here, the influence of an electric discharge—modeled by a heat-source term in the energy equation—on laminar boundary-layer flows over a flat plate with zero pressure gradient at Mach 3, 5, and 7 is investigated numerically. The aim was to appraise the potential of electro-gasdynamic devices for an application as turbulence generators in the super- and hypersonic flow regime. The results with localized heat-source elements in boundary layers are compared to cases with roughness elements serving as classical passive trips. The numerical simulations are performed using the commercial code ANSYS FLUENT (by ITAM) and the high-order finite-difference DNS code NS3D (by IAG), the latter allowing for the detailed analysis of laminar flow instability. For the investigated setups with steady heating, transition to turbulence is not observed, due to the Reynolds-number lowering effect of heating.
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react-text: 373 The results of numerical modeling of steady viscous compressible flow around The Earth descent module for Phobos-grunt interplanetary mission at Mach numbers 10 and 3 are presented. /react-text react-text: 374 /react-text
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react-text: 327 MBDA is leading an important research and technology effort on aero-propulsion covering a large set of propulsion technologies from high-speed dual-mode ramjet to detonation wave propulsion and hybrid electrical/chemical propulsion. MBDA is also working on the key issue of thermal management and developed innovative approach to control external flow thanks to near surface electrical discharges. /react-text react-text: 328 /react-text
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I. IntroductionIntensive studies of the physical processes of hypersonic boundary-layer tripping have been motivated by the needs of scramjet systems for air-breathing vehicles. For airframe-integrated scramjet engines, the forebody ahead of the inlet is designed to process and pre-condition the flow that will be ingested by the air inlet. Turbulent flow is desirable at the entrance to the inlet to mitigate flow separations on compression ramps and prevent air inlet unstarts [1-2]. However, natural transition typically does not occur on small-scale systems flying at high altitude (low Re number), thereby requiring boundary layer trip devices to ensure a turbulent boundary layer at the inlet. This issue motivated development of various boundary layer trip methods to promote turbulent flow in order to properly scale the engine flight test results to future full-scale vehicles. It was suggested that the most effective tripping mechanism requires the formation of streamwise vorticity within the boundary layer [3-6]. Typically, obstacles, moving elements, or non-steady gas jets are used to promote transition on the forebody. Recently, a thermal type of BL management was considered to be feasible [7].The non-thermal trips previously mentioned produce stationary forcing of the boundary-layer flow. They enhance and/or trigger instability mechanisms. However, they do not generate unstable disturbances. The latter are produced by the freestream noise or the incoming boundary layer. It is reasonable to assume that unsteady forcing, which generates disturbances of required length-scale and frequency, may be much more effective. Such an unsteady forcing can be produced by synthetic jets and/or plasma actuators.The primary mechanisms of plasma flow control include thermal (heating of the gas), electro-hydro-dynamic (EHD), and magneto-hydro-dynamic (MHD) interactions. EHD and MHD interactions induce bulk fluid motions via collisional momentum transfer from charged species accelerated in the plasma by Coulomb and Lorentz forces,
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This study was performed to characterize the dominant frequencies present in the boundary layer uptsream of and in the corner separation zone of a compression surface in Mach 4.5 flow and to determine a control effect of transient plasma actuation on the boundary layer. Schlieren imaging was used to distinguish the corner separation zone for 20°, 25°, and 30° compression ramps mounted on flat... [Show full abstract]
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A physical and numerical model for surface dielectric barrier discharge evolution in atmospheric air was developed and tested against experimental data for discharge parameters. Both discharge formation and relaxation phases were simulated successfully using a new approach of non-local air ionization by electron impact and ab initio boundary conditions on the electrode and dielectric surface. The main features of the physical and numerical model of the discharge simulation have been discussed. It was shown that discharge relaxation phase contributes primarily to momentum and heat sources relevant for flow control. The momentum source spatial distribution has a complex structure with the regions of upstream and downstream body force direction and qualitatively depends on applied voltage polarity and voltage pulse waveform. For different conditions it could lead to either near-surface flow acceleration or vortex generation.
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The development of a surface barrier discharge in air at atmospheric pressure under the action of a constant voltage of different polarity is simulated numerically. When the polarity of the high-voltage electrode is negative, the discharge develops as an ionization wave that moves along the dielectric surface. When the polarity is positive, the discharge develops as a streamer that first moves above the dielectric surface and then comes into contact with and continues to develop along it. In the case of a high-voltage electrode of positive polarity, the discharge zone above the dielectric surface is approximately five times thicker than that in the case of negative polarity. The characteristic aspects of numerical simulation of the streamer phase of a surface barrier discharge are discussed. The numerical results on the density of the charge stored at the dielectric surface and on the length of the discharge zone agree with the experimental data.
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Surface dielectric barrier discharge, initiated by a high-voltage pulse of negative polarity in atmospheric pressure air, is studied numerically and experimentally. At a pulse duration of a few tens of nanoseconds, two waves of optical emission propagate from the high-voltage electrode corresponding to the leading and trailing edges of the high-voltage pulse. It is shown by means of numerical modeling that a glow-like discharge slides along the surface of the dielectric at the leading edge of the pulse, slowing down on the plateau of the pulse. When the trailing edge of the pulse arrives to the high-voltage electrode, a second discharge starts and propagates in the same direction. The difference is that the discharge corresponding to the trailing edge is not diffuse and demonstrates a well-pronounced streamer-like shape. The 2D (in numerical modeling) streamer propagates above the dielectric surface, leaving a gap of about 0.05 mm between the streamer and the surface. The calculated and experimentally measured emission picture, waveform of the electrical current, and deposited energy, qualitatively coincide. The sensitivity of the numerical solution to unknown physical parameters of the model is discussed.
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A summary of recent research progress in hypersonic plasma actuators for flow control is attempted. It is found that the most effective plasma actuator is derived from an electromagnetic perturbation and amplifies by a subsequent viscous–inviscid interaction. Computational efforts using drift-diffusion theory and a simple phenomenological plasma model, as well as experiments in a hypersonic plasma channel, have shown the effectiveness of using electro–aerodynamic interaction as a hypersonic flow control mechanism. In principle, the plasma actuator based on magneto–aerodynamic interaction should have an added mechanism in the Lorentz force, making it even more effective as a flow control mechanism. However, this approach also incurs additional challenges and complications due to the Hall effect. Magneto–aerodynamic interactions have also been demonstrated for separated flow control, albeit in a very limited scope. Numerical simulations based on a simple phenomenological plasma model have shown the feasibility of separated flow suppression in shock-boundary-layer interaction over a compression ramp at a hypersonic flow of Mach 14.1. The control mechanism relies on the Lorentz force to energize the retarded shear layer in the viscous interacting region, but the effectiveness of momentum transfer via inelastic collision requires further validation.
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Plasma-fluid-dynamic interaction has been shown to be a viable mechanism for hypersonic flow control. An effective and verified flow control process using direct current surface discharge is summarized. The operating principle is based on a small electromagnetic perturbation to the growth rate of the displacement thickness of a shear layer that is strongly amplified by a subsequent pressure interaction. The aerodynamic control is delivered in less than a millisecond time frame and produces no parasitic effect when deactivated. The magnitude of the resultant aerodynamic force and moment can be significant and does not require a large amount of power for plasma generation to overcome the inefficient ionizing process, thus reducing the weight of a high-speed vehicle. The electromagnetic perturbation is derived from a surface gas discharge with or without an externally applied magnetic field. An embedded plasma actuator near the leading edge of a flat plate has produced high surface pressure equivalent to more than a 5 deg flow deflection at Mach 5, and the flow control effectiveness will increase with an increasing oncoming Mach number. The detailed flow structure of weakly ionized airstreams has been investigated by a combination of experimental effort and computational simulation solving the magneto-fluid-dynamic equations in the low magnetic Reynolds number limit with a drift-diffusion plasma model. The identical plasma actuator is investigated as a variable geometry cowl of a hypersonic inlet. All phenomena are replicated by computational results and are fully validated by experimental observations.
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To simulate the coupled plasma and fluid flow physics of dielectric-barrier discharge, a plasma luid model is utilized in conjunction with a compressible flow solver. The flow solver is responsible for determining the bulk flow kinetics of dominant neutral background species including mole fractions, gas temperature, pressure and velocity. The plasma solver determines the kinetics and energetics of the plasma species and accounts for finite rate chemistry. In order to achieve maximum reliability and best performance, we have utilized state-of-the-art numerical and theoretical approaches for the simulation of DBD plasma actuators. In this respect, to obtain a stable and accurate solution method, we tested and compared different existing numerical procedures, including operator-splitting algorithm, super-time-stepping, and solution of the Poisson and transport equations in a semi-implicit manner. The implementation of the model is conducted in OpenFOAM. Four numerical test cases are considered in order to validate the solvers and to investigate the drawbacks/benefits of the solution approaches. The test problems include single DBD actuator driven by positive, negative and sinusoidal voltage waveforms, similar to the ones that could be found in literature. The accuracy of the results strongly depends to the choice of time step, grid size and discretization scheme. The results indicate that the super-time-stepping treatment improves the computational efficiency in comparison to explicit schemes. However, the semi-implicit treatment of the Poisson and transport equations showed better performance compared to the other tested approaches.
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Nanosecond-pulsed surface dielectric barrier discharge (NS-DBD) plasma actuations with powered electrodes of different surface geometries were investigated numerically by solving the coupled plasma discharge equations, electron energy equations and the Navier-Stokes equations in quiescent air at atmospheric pressure. The plasma discharge characteristics and the air flow features were simulated numerically using a simple chemical kinetics plasma model for three powered electrodes with serrated, rectangular and semicircular surfaces, respectively. The results show that the reduced electric field of the serrated electrode is globally the strongest, while that of the rectangular electrode the second strongest, and that of the semicircular electrode the weakest. The maximum values of the reduced electric field, the mean electron energy and the electron density are found to occur immediately near the right upper tips of the powered electrodes, and the streamers of the mean electron energy and electron density in the serrated electrode case are larger in size and higher in value than in the rectangular and semicircular electrode cases. On the other hand, the pressure wave in the serrated electrode case is more intensive, and propagates slightly faster than in the other two electrode cases. Besides, the heated region in the serrated electrode case is greater with a higher temperature than in the other two electrode cases. The comparison results indicate that the performance of NS-DBD plasma actuators depends significantly on the powered electrode surface geometry, and the serrated surface design is a very promising means of flow control.
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A simplified (7 species and 9 processes) plasma kinetic model is proposed to investigate the mechanism of the plasma aerodynamic actuation driven by nanosecond-pulsed dielectric barrier discharge (NS-DBD). The governing equations include conservation equations for each species, the Poisson equation for the electric potential, and Navier-Stokes equations for the gas dynamic flow. Numerical simulations of plasma discharge and flow actuation on NS-DBD plasma actuators have been carried out. Key discharge characteristics and the responses of the quiescent air were reproduced and compared to those obtained in experiments and numerical simulations. Results demonstrate that the reduced plasma kinetic model is able to capture the dominant species and reactions to predict the actuation in complicated hydrodynamics. For the one-dimensional planar and two-dimensional symmetric NS-DBD, the forming of the sheath collapse is mainly due to the charge accumulation and secondary emission from the grounded electrode. Rapid species number density rise and electric field drop occur at the edge of the plasma sheath, where the space charge density gradient peaks. For the aerodynamic actuation with typical asymmetry electrodes, discharge characteristics have a core area on the right edge of the upper electrode, where the value can be much higher. The formation and propagation of the compression waves generated through rapid heating have also been performed and compared to those measured in a recent experiment. Energy release leads to gas expansion and forms a cylindrical shock wave, centering at the upper electrode tip with low gas acceleration. For the present single pulsed 1265kV case, the mean temperature of gas heating reaches about 57565K at 1μs and decreases to about 46065K at 10μs.
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Nanosecond pulse discharge plasma imaging, coupled pulse energy measurements, and kinetic modeling are used to analyze the mechanism of energy coupling in high repetition rate, spatially uniform, nanosecond pulse discharges in air in plane-to-plane geometry. Under these conditions, coupled pulse energy scales nearly linearly with pressure (number density), with energy coupled per molecule being nearly constant, in good agreement with the kinetic model predictions. In spite of high-peak reduced electric field reached before breakdown, E/N 500-700 Td, the reduced electric field in the plasma after breakdown is much lower, E/N 50-100 Td, predicting that a significant fraction of energy coupled to the air plasma, up to 30-40%, is loaded into nitrogen vibrational mode.
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Abstract Several components of a numerical procedure to simulate magnetoaerodynamic flow control are described and extended. A new element addresses the determination of the three-dimensional electric field for the general case where the conductivity is a spatially varying tensor due to the presence of Hall and ion-slip effects. The highly accurate discretization scheme employed to solve the flow equations in prior efforts is adapted to obtain the potential by solving the Poisson equation arising from current continuity. Boundary conditions for electrodes and insulators are derived for implementation on general 3-D curvilinear meshes. Several computations are utilized to verify the formulation, including a test case with an analytic solution and several cases for the potential arising between continuous and segmented electrodes with variable conductivity and velocity profiles. Issues related to the computation of the electric field in the high-speed regime are addressed by computing the initial induced field in a Type IV shock-on-shock interaction under a frozen velocity field assumption.
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The Navier Stokes equations were solved using a high-fidelity time-implicit numerical scheme and an implicit large-eddy simulation approach to investigate plasma-based flow control for supersonic flow over a compression ramp. The configuration included a flat-plate region to develop an equilibrium turbulent boundary layer at Mach 2.25, which was validated against a set of experimental measurements. The fully turbulent boundary-layer flow traveled over a 24 deg ramp and produced an unsteady shock-induced separation. A control strategy to suppress the separation through a magnetically-driven surface-discharge actuator was explored. The size, strength, and placement of the model actuator were based on recent experiments at the Princeton University Applied Physics Group. Three control scenarios were examined: steady control, pulsing with a 50% duty cycle, and a case with significant Joule heating. The control mechanism was very effective at reducing the time-mean separation length for all three cases. The steady control case was the most effective, with a reduction in the separation length of more than 75%. The controller was also found to significantly reduce the low-frequency content of the turbulent kinetic energy spectra within the separated region and reduce the total turbulent kinetic energy downstream of reattachment.
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The three dimensional phenomena in the weakly ionized plasma near an anode and a cathode of a Faraday type MHD generator are studied by time dependent three dimensional numerical analyses, where the radiation is taken into account. The Navier tokes equations are solved with an implicit total variation diminishing scheme with the radiative heat transfer solved by a finite volume method, while the Maxwell equations are solved by using the Galerkin finite element method. The following results are obtained. In the case of the short circuit load ( I=160 A), the strong Hall effect induces a strong electrical current concentration at the upstream edge of the anode and the downstream edge of the cathode, resulting in high temperature and high conductivity there. The Lorentz force acting at the spot of electrical current concentration may induce inter-anode breakdown, although the present generator does not suffer from inter-anode breakdown. The Lorentz force working at the electrical current concentration at the cathode, on the other hand, brings the concentration current downstream, resulting in strong inter-cathode breakdown. The radiative heat transfer becomes very high locally, resulting in 700 MW/m 3, but the effect can be neglected on the overall generator performance because the plasma is very dense.
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A conservative TVD scheme is adopted to solve the equations governing the three-dimensional flow of a nonideal compressible conducting fluid in a magnetic field. The eight-wave equations for magnetohydrodynamics (MHD) are proved to be a non-strict hyperbolic system, therefore it is difficult to develop its eigenstructure. Powell developed a new set of equations which cannot be numerically simulated by conservative TVD scheme directly due to its non-conservative form. A conservative TVD scheme augmented with a new set of eigenvectors is proposed in the paper. To validate this scheme, 1-D MHD shock tube, unsteady MHD Rayleigh problem and steady MHD Hartmann problem for different flow conditions are simulated. The simulated results are in good agreement with the existing analytical results. So this scheme can be used to effectively simulate high-conductivity fluids such as cosmic MHD problem and hypersonic MHD flow over a blunt body, etc.
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针对全 MHD(Magnetohydrodynamics)数值模拟中存在伪磁场散度的问题,发展了如下计算方法:基本格式基于八波对称形式方程组,补充相关源 项以保持方程组守恒性,并采用投影方法辅助清除伪散度.投影方法中,基于有限体积方法求解三维Poisson方程.算例显示,对于光滑解析磁场,伪磁场散 度得到有效清除;对于带激波高超声速MHD流动,全局投影下自由来流区域误差增大.提出一种局部投影方法,在高磁场散度区域进行投影.结果表明,最终流场 收敛稳定,高磁场散度得到有效清除,而低散度区域散度不受影响.
DOIURL [本文引用: 1]
针对全 MHD(Magnetohydrodynamics)数值模拟中存在伪磁场散度的问题,发展了如下计算方法:基本格式基于八波对称形式方程组,补充相关源 项以保持方程组守恒性,并采用投影方法辅助清除伪散度.投影方法中,基于有限体积方法求解三维Poisson方程.算例显示,对于光滑解析磁场,伪磁场散 度得到有效清除;对于带激波高超声速MHD流动,全局投影下自由来流区域误差增大.提出一种局部投影方法,在高磁场散度区域进行投影.结果表明,最终流场 收敛稳定,高磁场散度得到有效清除,而低散度区域散度不受影响.
