一种乘波前体进气道的一体化设计及性能分析

采用特征线方法设计了具有直线初始激波、内收缩段消除激波反射、出口参数均匀可控的基准内锥流场.基于密切内锥(osculating inward turning cone,OIC)乘波体设计方法,发展了一体化密切内锥乘波前体进气道(osculating inward turning cone waverider inlet,OICWI)设计技术.基于一体化基准内锥流场和前体进气道设计技术,设计了密切内锥乘波前体进气道.采用数值软件对设计的乘波前体进气道进行了仿真分析,结论如下:①OICWI的设计是遵循气动原理的.②一体化密切内锥乘波前体进气道的前缘形状、内收缩比及出口参数可以根据需求定量准确设计.③理论设计结果和模拟结果吻合一致,证明设计方法是正确可靠的.④数值模拟研究结果表明一体化密切内锥乘波前体进气道具有较好的出口流场均匀度及较高的流量捕获率和较高的总压恢复特性.      

展向截断曲面乘波压缩进气道气动布局

描述了所设计的展向截断曲面乘波压缩进气道.其特点是采用曲面乘波压缩前体,前体进气道压缩面基准流场由等熵压缩波轴对称流场组成,三维乘波面采用密切曲锥方法由前缘线各点流线跟踪拟合构成流面.乘波面根据飞行器和发动机的宽度要求进行了截断.数值计算和风洞试验结果表明:与相同几何收缩比的四波系压缩进气道相比,在马赫数为4.5时,曲面乘波压缩进气道流量系数提高12%,总压恢复系数提高39%;在马赫数为6时,曲面乘波压缩进气道流量系数提高4%,总压恢复系数提高50%.超然冲压发动机性能明显提高.      

低反动度高负荷超声速轴流压气机气动设计方法

对超声速轴流压气机的发展历程进行了总结,分析了两类传统超声速压气机内部流动特点及其所存在的流动问题,并对未来超声速压气机的发展提出了展望.针对超声速压气机内部流动特点,提出了一种新的低反动度高负荷超声速轴流压气机气动设计原理.该原理可有效避免在动叶中进行流动控制,同时降低动叶出口绝对马赫数,并结合附面层抽吸控制静叶栅内部流动,最终实现级的高负荷气动设计.利用该原理进行设计验证,三维黏性数值模拟结果表明:在叶尖切线速度为360m/s的前提下,实现了2.3的级压比,级效率为86.5%.      

无激波机翼的设计方法

 本文介绍了用虚拟气体概念进行无激波机翼设计的方法,叙述了虚拟气体方法的基本思路;给出了几种虚拟气体规律,以小扰动方程为例具体说明了在超音区内的推进计算过程。分别讨论了二维和三维情况,设计和非设计状态时的效益。最后介绍了近无激波机翼修型的工程设计方法。     

超声速逆向喷流的减阻与降热

利用流动控制技术改变流场结构的方法,对逆向冷喷流将激波推离物体表面以达到减阻减热的效果,进行了数值模拟计算。通过计算给出了钝头体在超声速流场中有无逆向喷流的压力场和温度场的比较,结果表明,逆向喷流对钝头体表面的减阻和减热起到了明显的作用。     

亚超音速机翼最佳弯扭综合设计计算方法

本文介绍一种亚超音速机翼最佳弯扭综合设计的计算方法,它应用了有限基本解方法。分别在亚超音速各选取一个设计点(M数和C_L),进行机翼弯扭设计,其目的是减小与升力相关的阻力。在此基础上,顾及亚音速和超音速这两个设计点的气动力特性,还要兼顾到飞机其它性能和结构上实现的可能性,进行机翼的综合设计。本文分别给出了亚音速最佳弯扭设计,超音速最佳弯扭设计和综合设计的计算结果。经过分析表明,计算结果是合理的。     

飞行器BLISS分布式平台的研究与实现

用matlab和VC 作为开发工具,开发出基于tcp/ip协议和windows socket的适合于飞行器BLISS的软件BLISS Platform。BLISS Platform采用客户端/服务器构架,集成了协作优化算法,是适合于多学科设计优化的分布式计算机网络软件平台。平台的底层通讯模块由VC 程序实现,仿真优化程序由matlab开发,在matlab环境下修改飞行器数学模型就可实现对不同飞行器模型的优化设计。最后用BLISS Platform并行实现了一简单的超音速飞行器概念模型的优化设计,取得了较好的结果。     

超声速低反力度吸附式压气机变工况特性

为探讨非设计点的气动性能,以某3级高负荷低反力度吸附式压气机首级超声速级作为研究对象,借助数值模拟的方法,分别研究了变转速、变抽吸量以及近失速点首级超声速级参数以及内部流场的变化.结果表明:变转速条件下,基于低反力度设计概念研制的吸附式压气机所采用的抽吸方案仍能适应设计要求;减少抽吸量,在降低本级气动性能的同时,增加了其与下一级流动匹配的难度;通过在端壁抽吸槽道前段添加附加抽吸槽道/孔可进一步拓宽吸附式压气机的工作范围.      

液晶流动可视化方法研究拟似冲击波的内部超声速流

在一个压力一真空超声速风洞中,剪切应力敏感液晶流动可视化技术被应用来研究方管内马赫数2拟似冲击波（pseudo-shockwave）的超声速流动。它主要提供关于整个流动的定性信息,诸如湍流边界层分离、再附着位置以及流动的维数等。而且液晶也反映了表面流线,分离区内的涡流和管道流动的角效应。使用两种不同黏度的液晶分别进行实验,分析黏度对结果的影响。液晶实验的结果与纹影照片所得结果比较吻合,说明了液晶是一个非常有效的流动可视化工具。     

槽壁试验段低超声速流场特性数值模拟

 在跨声速风洞中通过开槽和抽气可以建立低超声速流场,由于槽壁试验段设计参数多,流场结构复杂,为提高设计准确性,通过数值模拟对槽壁试验段低超声速流场特性进行了研究。首先根据槽壁试验段的一般设计准则进行了气动设计,给出了槽壁尺寸和不同马赫数所对应的抽气量。基于设计结果,通过数值模拟对流场特性进行了研究,计算表明：通过抽气可以建立均匀的低超声速流场,抽气量对试验段马赫数均匀区长度有较大影响;随后对不同气动外形进行了比对,结果表明:抽气口位置、壁板厚度、驻室容积、开闭比及槽壁外形等对试验段的气流质量有影响,开闭比和槽型的影响尤为显著;最后对槽壁的通流特性进行了分析,探讨了槽型对试验段流场影响的成因。     

基于密切原理的Bump进气道外压缩鼓包逆向设计

基于密切乘波理论提出一种Bump进气道外压缩鼓包的设计方法,可根据制定的激波形状及其曲率中心分布来逆向求解外压缩鼓包型面。通过引入曲率中心分布这一变量,可以控制横截面激波形状并调节外压缩鼓包的三维外型及其表面横向压力分布,进而提高外压缩鼓包的附面层排移能力。同时,发展了一种Bump进气道的流量系数快速估算法,能够在设计初期以不超过2%的误差快速给出进气道的流量系数。结果表明:基于密切原理的外压缩鼓包设计有利于改进Bump进气道的流量捕获和附面层排移能力。算例中,较锥导鼓包模型,密切鼓包方法设计的Bump进气道流量系数提升4.03%,附面层排移能力提升2.12%。      

基于超声速有益干扰原理的气动构型概念综述

超声速有益干扰气动设计概念于20世纪30年代提出，其基本思想是利用飞行器部件间的波系干扰获得诸如增升或减阻等性能收益。此概念在20世纪50~60年代得到了大量探索并部分实现了工程应用，在20世纪70年代至世纪末陷入沉寂。近年来，随着超声速运输机和高超声速飞行器技术的复兴，超声速有益干扰概念重新得到重视并有望得到工程应用。本文梳理了超声速有益干扰气动设计概念的发展历史，总结了应用超声速有益干扰原理的典型构型，如超声速双翼机、Flat-top构型、环翼和半环翼构型、伞翼构型、高压捕获翼构型等，并对典型构型的基本原理和气动特点进行了分析。对超声速有益干扰设计概念的未来进行了展望，概述了亟待研究的相关问题。     

吻切锥乘波构型参数化设计与正交试验分析

为了获得外形控制参数对吻切锥乘波构型性能的影响程度,通过分析吻切锥乘波构型的生成特点和生成方法,在其参数化设计的基础上,运用正交试验设计方法分析了各个控制参数乘波构型性能的影响,确定了对气动性能和容积率影响较大的参数,为进一步合理确定优化空间和优化策略提供指导。运用CFD方法对典型外形进行了性能分析,结果显示:吻切锥乘波构型具有较高升阻比,下表面中心区流动均匀,可为高超声速飞行器机身/进气道一体化提供参考。作为应用,基于正交试验结果设计了以吻切锥乘波构型作为前体的一个高超声速飞行器,验证了设计方法的合理性。     

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.     

台阶后横喷氢气超音速燃烧流场数值模拟研究

采用ＭｃＣｏｒｍａｃｋ格式、代数涡粘性湍流模型及有限速率化学动力学模型，用数值模拟方法研究了台阶后横喷氢气二元燃烧流场。数值计算结果表明：台阶的作用不仅能扩大火焰稳定性，而且增加氢射流对主流的穿透深度，提高燃烧效率；当进口气流Ｍ数越高，进口温度越高，油气当量比越接近于恰当比，壁面温度对流场的影响越大。还提出了在超音速气流中，横喷氢自动着火时滞的火焰稳定机理新观点，由此可更准确地预估自动着火点的位置及火焰稳定性。     

新型斜波瓣超声速混合器设计研究及其数值计算

根据超声速两股流掺混的特点, 在波瓣喷管的基础上提出了一种新的斜波瓣型超声速混合器设计方法, 设计的原则是在较小流动损失的前提下尽可能增加混合程度.利用有限体积法对三维可压粘流N-S方程进行离散, 紊流模型为k-ε模型, 分别对欠膨胀、接近完全膨胀和过膨胀3种工况进行了计算研究.计算结果表明, 当内通道瓣腔顶端与外通道相交处成一定夹角, 射流产生鞍形激波和旋涡可以增加掺混效率, 混合器内两股超声速流的流态基本符合设计要求.在欠膨胀、接近完全膨胀的工况, 效果是令人满意的, 但对严重过膨胀的工况, 到出口截面的混合效果不理想.      

各向异性网格自适应计算在超燃模拟中的应用

采用各向异性网格自适应求解技术,将其应用于DLR超燃冲压发动机燃烧室中的超燃模拟.开展了3个算例,包括采用滑移壁面条件的燃烧室冷流场模拟、采用无滑移壁面条件的冷流场模拟及采用无滑移壁面条件的反应流模拟.模拟中,各向异性网格自适应计算捕捉到了如激波、射流、边界层、火焰面等具有各向异性特征的大梯度区域,并利用各向异性网格进行了很好的加密.对比利用各向同性网格的初始流场计算,各向异性网格自适应计算使基于滑移条件、无滑移条件的冷流计算及反应流计算的网格单元数量分别下降了36.2%,36.4%和36.8%,有效降低了计算规模,而且流场大梯度区域的计算结果更准确,辨析度更好.结果表明:对于像超燃这类具有各向异性特征的问题,各向异性网格系统比各向同性网格系统有更好的计算效率及准确性,同时也表明基于Mach数场构造的各向异性网格系统可以有效应用于超燃计算.      

Supersonic biplane—A review

One of the fundamental problems preventing commercial transport aircraft from supersonic flight is the generation of strong sonic booms. Sonic booms are the ground-level manifestation of shock waves created by airplanes flying at supersonic speeds. The strength of the shock waves generated by an aircraft flying at supersonic speed is a direct function of both the aircraft’s weight and its occupying volume; it has been very difficult to sufficiently reduce the shock waves generated by the heavier and larger conventional supersonic transport (SST) configuration to meet acceptable at-ground sonic-boom levels. It is our dream to develop a quiet SST aircraft that can carry more than 100 passengers while meeting acceptable at-ground sonic-boom levels. We have started a supersonic-biplane project at Tohoku University since 2004. We meet the challenge of quiet SST flight by extending the classic two-dimensional (2-D) Busemann biplane concept to a 3-D supersonic-biplane wing that effectively reduces the shock waves generated by the aircraft. A lifted airfoil at supersonic speeds, in general, generates shock waves (therefore, wave drag) through two fundamentally different mechanisms. One is due to the airfoil’s lift, and the other is due to its thickness. Multi-airfoil configurations can reduce wave drag by redistributing the system’s total lift among the individual airfoil elements, knowing that wave drag of an airfoil element is proportional to the square of its lift. Likewise, the wave drag due to airfoil thickness can also be nearly eliminated using the Busemann biplane concept, which promotes favorable wave interactions between two neighboring airfoil elements. One of the main objectives of our supersonic-biplane study is, with the help of modern computational fluid dynamics (CFD) tools, to find biplane configurations that simultaneously exhibit both traits. We first re-analyzed using CFD tools, the classic Busemann biplane configurations to understand its basic wave-cancellation concept. We then designed a 2-D supersonic biplane that exhibits both wave-reduction and cancellation effects simultaneously, utilizing an inverse-design method. The designed supersonic biplane not only showed the desired aerodynamic characteristics at its design condition but also outperformed a zero-thickness flat-plate airfoil. (Zero-thickness flat-plate airfoils are known as the most efficient monoplane airfoil at supersonic speeds.) Also discussed in this paper is how to design 2-D biplanes, not only at their design Mach numbers but also at off-design conditions. Supersonic biplanes have unacceptable characteristics at their off-design conditions such as flow choking and its related hysteresis problems. Flow choking causes rapid increase of wave drag and it continues to be kept up to the Mach numbers greater the cruise (design) Mach numbers due to its hysteresis. Some wing devices such as slats and flaps, which could be used at take-off and landing conditions as high-lift devices, were utilized to overcome these off-design problems. Then supersonic-biplane airfoils were extended to 3-D wings. Because that rectangular-shaped 3-D biplane wings showed undesirable aerodynamic characteristics at their wingtips, a tapered-wing planform was chosen for the study. A 3-D biplane wing having a taper ratio and aspect ratio of 0.25 and 5.12, respectively, was designed utilizing the inverse-design method. Aerodynamic characteristics of the designed biplane wing were further improved by using winglets at its wingtips. Flow choking and its hysteresis problems, however, occurred at their off-design conditions. It was shown that these off-design problems could also be resolved by utilizing slats and flaps. Finally, a study on the aerodynamic characteristics of wing-body configurations was conducted using the tapered biplane wing. In this study a body was chosen in order to generate strong shock waves at its nose region. Preliminary parametric studies on the interference effects between the body and the tapered biplane wing were performed by choosing several different wing locations on the body. From this study, it can be concluded that the aerodynamic characteristics of the tapered biplane wing are minimally affected by the disturbances generated from the body, and that the biplane wing shows promise for quiet commercial supersonic transport.     

NPLS技术及其在高速飞行器气动研究中的应用

近年来,与高速飞行器相关的超声速/高超声速流动受到了极大关注。这类流动所具有的非定常性、强梯度和可压缩性对试验研究提出了挑战。纳米示踪的平面激光散射技术（NPLS）是2005年由作者所在的研究团队研发的非接触光学测试技术。它能够获得超声速三维流场的某个剖面的瞬态流动结构,并且具有较高的时空分辨率。目前,许多研究结果表明NPLS是研究超声速湍流的一项非常有效的技术。近年来,作者应用 NPLS 技术在超声速湍流研究中取得了较大的进展,并且基于NPLS开发了其它几种技术,比如基于 NPLS 的密度场测量技术（NPLS-DT）,能够获得超声速流动的密度场信息并还能进一步得到雷诺应力分布。本文介绍了NPLS技术并回顾了其在超声速边界层、激波/边界层相互作用等流动中的应用。由于能够获得雷诺压力和湍动能等统计量, NPLS技术有望在发展可压缩湍流模型的研究中发挥作用。     

一种二元定几何混压式超声速进气道流场控制概念研究

针对二元定几何混压式超声速进气道低马赫数时流量系数低加速性能差的问题,提出了一种新的泄流槽流场控制概念,并通过数值仿真,揭示了泄流槽控制激波结构机理及其主要几何参数对进气道性能的影响规律.研究结果表明:采用该流场控制方案可通过泄流槽入口处的波系结构使进气道在低于设计马赫数时的出口总压恢复系数和流量系数相对于原型方案均得到明显提高,而在设计点关闭泄流槽后进气道的性能与原型进气道基本相当,这对改善冲压发动机在低马赫数转级后的加速性能是有利的.