基于共轭方程法的跨音速机翼气动力优化设计

 设计状态的机翼气动力特性是设计人员最为关心的指标, 应用控制理论设计方法进行了有升力约束情形下跨音速机翼阻力优化设计研究, 根据给定的目标函数推导了相应的共轭方程和边界条件, 研究了共轭方程的数值求解方法, 以及计算目标函数对设计变量的敏感性导数时所涉及的度量矩阵变分求解问题, 研究了流场计算、共轭方程数值求解、敏感性导数求解和拟牛顿优化算法这几个主要方面的有效结合问题, 发展出了一种跨音速机翼气动力优化设计方法, 进行了跨音速机翼气动力优化设计研究验证, 优化后机翼气动力特性有一定程度的改善, 阻力系数能减少20%左右, 而升力系数有所增大, 说明所发展的设计方法是成功的, 该设计方法在跨音速及复杂外形气动设计方面比以往设计方法具有更好的适用性和优越性。     

发动机进排气系统对飞机气动特性影响的数值研究

为了分析发动机与飞机系统耦合过程中产生的气动力特性,准确获取实际使用条件下的飞机升阻特性,以带动力的某半模飞机模型为研究对象,开展了三维流场数值仿真计算,利用风洞试验数据对计算方法进行了验证,总结了溢流阻力及排气干扰阻力的变化规律,结果表明:发动机状态变化对飞机升阻特性影响明显,在选定的发动机工作参考状态下,溢流阻力在飞机系统升力与阻力方向的分量随捕获面积比的增大分别增大、减小,排气系统实际工作状态偏离选定的参考状态越远,干扰阻力在飞机系统升力和阻力方向分量的绝对值越大。     

用计算流体力学解读急剧机翼失速

文章描述了在为期4年的急剧机翼失速（AWS）国家研究计划中用计算流体动力学所做的工作及获得的经验；详细论述了在跨音速机动状态现代美国战斗机和强击机遭遇到的跨音速流的复杂特性，内容包括针对这种高度复杂、有激波诱导的大分离粘性流动的网格处理、计算存储器和处理器要求、湍流模型、定常和非定常计算以及高度复杂的雷诺平均NS方程解与脱体涡模拟方法的比较。对F／A-18C飞机、AV-8B飞机、预生产F／A-18E飞机和F-16C飞机的跨音速机动状态进行了CFD模拟。模拟了各种襟翼位置，并将计算结果与大规模的风洞数据进行了比较；将结构与非结构网格获得的结果进行了对比；通过对模型在风洞中获得的表面油流和压敏涂层（psp）试验结果的定性比较以及有试验结果的详细定量压力系数数据来评估不同CFD方程式的适用性和计算精度。对比了这些计算软件和在这次研究中涉及的每种飞机的试验数据之间的静升力系数。     

翼吊布局长涵道发动机短舱气动布局优化研究

采用CFD数值计算的方法,对翼吊发动机长涵道短舱相对于机翼的前伸量和下沉量进行优化分析.分析了发动机短舱不同前伸量、下沉量时,飞机在高速巡航状态的干扰阻力和升力损失,以及满足低速平飞最小可用升力下,大迎角状态的失速特性,并计算分析了马赫数对优化结果的影响.计算结果表明,前伸量X/C =0.1时,干扰阻力及升力损失较小;...     

跨音速三维多升力面气动设计

发展了一种跨音速多升力面的气动设计方法和设计程序,它基于已成功应用于亚、跨音速机翼设计和亚音速双翼面设计的"余量修正迭代"概念.当升力面上出现超音速区和激波时自动引用迎风格式对控制方程进行修正.开发了一系列接口程序,包括目标压力设计程序.由此气动设计程羊、TAU程序以及相应的接口程序建立了跨音速多升力面气动反设计软件系统.用两个鸭翼-机翼构型验证了设计方法和设计程序,结果表明在高跨音速下设计迭代有很好的收敛性.     

SA一方程湍流模型参数影响分析与辨识

SA(Spalart-Allmaras)一方程湍流模型是目前工程湍流计算中主要采用的湍流模型之一。首先,针对某典型战斗机的小攻角、中等攻角、大攻角流动工况,利用均匀试验设计方法分析SA一方程模型中8个参数取值对上述工况下飞机升力和阻力系数计算结果的影响规律;然后,建立工程湍流模型参数的辨识方法,并将其用于该战斗机大攻角工况下湍流模型参数cb1值的辨识调整。结果表明:不同工况下,湍流模型参数对计算结果的影响规律不同;在附着流状态下,对升力和阻力影响较大的参数是cv1和cb1;在中等攻角和较大攻角下,对升力和阻力影响较大的参数是cb1;适当减小参数cb1的取值后,升力和阻力系数的计算结果有较明显的改善,这可能与飞机大攻角分离流场中涡粘系数和剪切应力的发展与自由剪切流存在一定差异有关。     

降雨对飞机性能影响的计算分析

计算了在五种不同降雨率条件下飞行时飞机表面增加的雨水重量以及雨水施加给飞机的水平冲力和垂向冲力,分析了降雨对飞机阻力增量的影响和升力损失特性,研究了风切变条件下的降雨对飞机性能的影响。最后得出了降雨对飞机性能影响的几点结论。     

民用飞机高升力系统设计载荷计算

民用飞机高升力系统是影响飞机起降性能的一个关键系统,根据高升力系统设计特点,探讨翼面气动载荷、驱动装置驱动能力、设备传动效率和阻力、卡阻、脱开故障与系统设计载荷关系,系统阐述由设计输入襟缝翼气动载荷计算高升力系统各设备的设计载荷计算流程,得到适用于工程应用的高升力系统载荷计算方法.     

民机高速风洞试验的阻力雷诺数效应修正

在飞机设计阶段,获得准确的气动特性数据对设计者是十分重要的,采用缩比模型的风洞试验是主要途径之一。但是风洞条件限制了试验的Re数,从而影响了对跨音速巡航阻力的评估,因此需要对风洞试验结果的阻力进行Re数效应修正。分析了风洞试验结果阻力的误差构成,提出了Re数效应修正的风洞试验数据基础;进一步就跨音速巡航阻力的构成进行分析,论述了对最小阻力和升致阻力分项修正的方法,并对修正结果进行了验证,得到了一种可靠的工程方法。     

一种工程实用的跨音速翼型多点优化设计方法

将N-S方程流场分析程序和SQP（序列二次规划）结合起来，发展了一种工程实用的跨音速翼型优化设计方法，用以提高翼型在多个设计点、多种约束条件下的气动性能。由N-S方程计算得到的升力、阻力等气动参数构成目标函数，用数值优化程序对其进行最优化。超临界翼型的设计实例表明，该方法设计质量高，所需机时少，易于实施，有较大的工程应用价值。     

某运输机全机构型雷诺数效应阻力修正研究

采用CFD方法分析了雷诺数效应对某运输机全机构型阻力的影响。基于对某运输机的翼/身/架/舱组合体绕流进行的计算分析,归纳总结出了一种适应于复杂构型的低雷诺数到高雷诺数的阻力修正方法。通过验证可知,该修正方法修正效果良好,为工程中的阻力预测提供了一种简洁有效的方法。     

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。     

Aerodynamic Design Methodology for Blended Wing Body Transport

This paper puts forward a design idea for blended wing body(BWB).The idea is described as that cruise point,maximum lift to drag point and pitch trim point are in the same flight attitude.According to this design idea,design objectives and constraints are defined.By applying low and high fidelity aerodynamic analysis tools,BWB aerodynamic design methodology is established by the combination of optimization design and inverse design methods.High lift to drag ratio,pitch trim and acceptable buffet margin can be achieved by this design methodology.For 300-passenger BWB configuration based on static stability design,as compared with initial configuration,the maximum lift to drag ratio and pitch trim are achieved at cruise condition,zero lift pitching moment is positive,and buffet characteristics is well.Fuel burn of 300-passenger BWB configuration is also significantly reduced as compared with conventional civil transports.Because aerodynamic design is carried out under the constraints of BWB design requirements,the design configuration fulfills the demands for interior layout and provides a solid foundation for continuous work.     

民用客机可变弯度机翼优化设计研究

远程宽体客机在实际飞行状态下,机翼变弯度有效减阻能够提升客机性能和飞行品质。以全机配平构型为研究对象,基于襟翼、扰流板偏转建立变弯度模型;采用RANS方程实现阻力的精确求解并建立响应面模型,对不同升力系数、马赫数的多个飞行状态进行变弯度减阻优化;在此基础上,对实际飞行过程中变弯度操作需求及综合减阻性能进行分析,并采用布雷盖公式评估机翼变弯度后全航段综合巡航效率。结果表明：在巡航马赫数飞行时,采用两次变弯度设置即可在较宽的升力系数范围内获得减阻收益;在10 km定高巡航时,机翼变弯度可使整个航段综合减阻1.9 cts,航时、航程提高0.72%;在8和10 km进行一次阶梯巡航,机翼变弯度可使整个航段综合减阻2.9 cts,航时、航程提高1.19%。     

锥形流乘波体优化设计研究

随着马赫数的升高,波阻和摩阻增加,形成升阻比"屏障",而乘波构型飞行器是克服这一屏障的有效途径。本文在Ma_∞=4.0~20.0、高度H=24.0~52.0km、圆锥角Ac=5°~10°的条件下,以升阻比为目标函数,进行了乘波体的优化设计,讨论了对乘波体优化外形的影响因素,并给出全马赫数范围的优化乘波体外形及其气动力结果。     

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.     

全动翼尖对无尾飞翼布局飞机气动特性影响的实验研究

通过风洞实验对无尾飞翼布局飞机的本体气动特性和全动翼尖的操纵效能进行了研究。结果表明,此布局飞机是纵向静稳定的,但零升俯仰力矩系数为负值,这对飞机的起降不利。全动翼尖作动时将增加飞机的阻力,降低全机的最大升阻比,产生抬头力矩。全动翼尖单侧作动可在升力基本保持不变的条件下提供偏航力矩,但同时也伴随着较大的滚转力矩和俯仰力矩。此外,全动翼尖的单侧作动和侧力是弱耦合的。全动翼尖同步作动的纵向特性和单侧作动规律相似,但幅度相对要大。     

翼/短舱/吊架气动特性及其动力影响数值研究

应用计算流体动力学(CFD)数值模拟方法对DLR-F6模型分别在通流及动力条件下翼/短舱/吊架(W-N-P)组合体部分的气动干扰特性进行研究.结果表明:组合体相互干扰引起的干扰通道附增激波及短舱后部逆压导致的涡流增加了短舱外罩阻力,降低了机翼升力;动力条件较通流条件,F6短舱外罩阻力降低,机翼升阻力特性在零度攻角时有所提高.      

趋于临界马赫数的圆柱跨声速绕流特性分析

通过深化认识趋于临界马赫数Ma_cr的圆柱跨声速绕流特性,明确新型飞行器增升减阻设计的空气动力学理论依据。采用大涡模拟方法数值研究了来流马赫数Ma_∞为0.75和0.85、雷诺数Re为2×10⁵的圆柱跨声速绕流。结果表明:当Ma_∞趋于临界马赫数(Ma_cr≈0.9)时,圆柱的阻力下降且升力系数振荡被抑制;通过力的分解,得知圆柱的阻力减小来自旋涡力的影响,而非可压缩性;圆柱的阻力减小与其背压上升有关;剪切层初始阶段的对流马赫数Ma_c随Ma_∞的增加而增大,而增长率相反,这使得剪切层更为稳定、柱体背压更高。此外,由于Ma_∞=0.85时边界层分离点处的激波和尾迹处的激波向下游推移,使得近尾迹处的湍流脉动减弱,进而导致柱体的表面压力振荡和升力系数振荡被抑制。     

钝化前缘乘波布局及其一体化构型气动特性

以最大升阻比为优化目标,在锥型流场中优化设计出乘波布局,并考虑高超声速飞行器的防热需求,对乘波布局进行钝化设计,利用数值模拟和风洞实验两种手段,研究钝化前缘乘波布局的气动特性.结果表明:在一定钝化半径内,随着钝化半径的增加,乘波构型的升力特性变化仅为2%,但阻力特性增加近3倍,升阻比降低了将近50%.尽管如此,为了钝化乘波布局,仍维持了较高的升阻比,升阻比为3左右.同时,以二维顶压式进气道为基础,在多级楔锥组合体流场中,设计出满足超燃发动机进气要求的乘波前体/进气道一体化构型,并进行前缘钝化设计.针对一体化构型进行了数值验证,结果表明:此类一体化构型升阻比大于2.6,同时发动机总压恢系复数保持在40%左右,满足进气道的要求.