基于实验设计方法的高超声速飞机前缘型线优化分析

为探索前缘线变化对吸气式高超声速飞机气动性能的影响,基于一种旁侧进气布局翼身融合体构型,在飞行马赫数6,攻角4°和高度26 km的巡航飞行条件下,结合运用增量修正参数化设计方法、均匀实验设计方法和计算流体力学模拟,分析了飞行器前缘型线与其升阻力系数及纵向压心等性能参数间的关系.计算结果表明,前缘线形状对飞行器升阻力系数明显高于其对纵向压心影响,设计空间范围内升力系数变化约21.3%,阻力系数变化约31.8%,升阻比变化范围约10.63%,但相对压心变化范围仅为3.87%.在此基础上,通过对典型构型物面压力分布进行分析,发现前缘线形状适当弯曲可利用飞行器下表面侧壁压缩产生的高压气流,利用二者的耦合效应使飞行器获得额外的升力增量.      

高压捕获翼前缘型线优化和分析

为分析翼前缘形状变化对高压捕获翼构型气动性能的影响,基于一种锥体组合捕获翼概念构型,采用幂次函数和余弦函数组合形式对翼前缘型线进行了参数化设计,在比较了多项式和径向基函数两种代理模型的拟合精度基础上,以飞行马赫数7,飞行攻角0° 为计算条件,结合使用均匀实验设计方法、计算流体力学、径向基函数代理模型方法和遗传算法,选择升阻比最大化为目标开展了数值优化,最后基于优化结果进行了单参数的灵敏度分析. 优化结果表明,相对于基准外形而言,优化后构型升力系数增大了约8.1%,阻力系数减小了约12.2%,升阻比提高了约23.4%. 此外,灵敏度分析结果表明升阻比与5 个设计参数均呈非线性关系,其中展向角度对升阻比影响最大,其次为幂次曲线的比例参数,其余3 个参数对升阻比的影响相对较弱.      

Numerical evaluation of passive control of shock wave/boundary layer interaction on NACA0012 airfoil using jagged wall

Shock formation due to flow compressibility and its interaction with boundary layers has adverse effects on aerodynamic characteristics, such as drag increase and flow separation. The objective of this paper is to appraise the prac-ticability of weakening shock waves and, hence, reducing the wave drag in transonic flight regime using a two-dimensional jagged wall and thereby to gain an appropriate jagged wall shape for future empirical study. Different shapes of the jagged wall, including rectangular, circular, and triangular shapes, were employed. The numerical method was validated by experimental and numerical studies involving transonic flow over the NACA0012 airfoil, and the results presented here closely match previous experimental and numerical results. The impact of parameters, including shape and the length-to-spacing ratio of a jagged wall, was studied on aerodynamic forces and flow field. The results revealed that applying a jagged wall method on the upper surface of an airfoil changes the shock structure significantly and disinte-grates it, which in turn leads to a decrease in wave drag. It was also found that the maximum drag coefficient decrease of around 17%occurs with a triangular shape, while the max-imum increase in aerodynamic efficiency (lift-to-drag ratio) of around 10%happens with a rectangular shape at an angle of attack of 2.26?.     

高超声速钝前缘乘波构型优化设计研究

乘波体的高升阻比优势使其在高超声速飞行器设计中极具应用前景. 在实际工程应用中, 为了满足防热要求, 乘波体前缘必须进行钝化处理, 前缘钝化对乘波体气动性能会产生显著影响. 因此, 原始尖前缘最优乘波体并不能保证钝化后仍为最优. 针对这一问题, 首先研究了前缘钝化对不同构型升阻特性的影响程度和作用机理. 结果表明: 前缘钝化会造成乘波体升力小幅度降低, 阻力大幅增加, 升阻比显著降低; 其中钝前缘本身的波阻在阻力增量中起主导作用, 而钝前缘本身的摩阻增加量与物面的摩阻降低量十分接近. 基于上述结果, 提出了一种高效评估钝前缘乘波体气动力的方法, 并结合遗传算法, 开展了直接考虑前缘钝化影响的乘波构型优化设计研究, 获得了钝前缘最优构型. 通过CFD数值模拟对最优构型的气动力特性进行评估, 结果表明: 在不同飞行高度、不同升力和不同钝化半径约束下, 相比尖前缘最优构型, 钝前缘最优构型宽度变窄, 相同纵向位置处的后掠角增大, 且升阻比显著提升. 在M_∞ = 15, H = 50 km, CL = 0.3约束条件下, 钝化半径R = 10 mm的钝前缘最优构型设计点升阻比相比尖前缘最优构型提升量可达9.32%.      

高超声速飞行器气动防热新概念研究

传统乘波构型的高超声速飞行器尖锐的前缘存在严重的气动加热问题,而简单的前缘钝化气动防热方法由于造成很大的升阻比损失,难以发挥实质性作用. 引入``人工钝前缘(ABLE)'概念,拟以一种新的思路解决这一矛盾. 通过定义ABLE构型的外形参数,并采用CFD数值计算方法研究了各参数对气动力和气动热特性的影响规律,在流场分析的基础上进行了外形优化,最终得到令人满意的新型高超声速飞行器头部外形,总结了运用ABLE概念进行气动防热的相关设计原则和规律.      

The application of the gradient-based adjoint multi-point optimization of single and double shock control bumps for transonic airfoils

A shock control bump (SCB) is a flow control method that uses local small deformations in a flexible wing surface to considerably reduce the strength of shock waves and the resulting wave drag in transonic flows. Most of the reported research is devoted to optimization in a single flow condition. Here, we have used a multi-point adjoint optimization scheme to optimize shape and location of the SCB. Practically, this introduces transonic airfoils equipped with the SCB that are simultaneously optimized for different off-design transonic flight conditions. Here, we use this optimization algorithm to enhance and optimize the performance of SCBs in two benchmark airfoils, i.e., RAE-2822 and NACA-64-A010, over a wide range of off-design Mach numbers. All results are compared with the usual single-point optimization. We use numerical simulation of the turbulent viscous flow and a gradient-based adjoint algorithm to find the optimum location and shape of the SCB. We show that the application of SCBs may increase the aerodynamic performance of an RAE-2822 airfoil by 21.9 and by 22.8 % for a NACA-64-A010 airfoil compared to the no-bump design in a particular flight condition. We have also investigated the simultaneous usage of two bumps for the upper and the lower surfaces of the airfoil. This has resulted in a 26.1 % improvement for the RAE-2822 compared to the clean airfoil in one flight condition.     

涡波一体乘波飞行器宽速域气动优化设计研究

涡波一体宽速域乘波飞行器通过在低速引入涡效应,显著改善了传统乘波体在低速状态下的升阻特性,具有在未来宽速域空天飞行器总体气动设计当中得到广泛应用的巨大潜力.但是,该设计方法的研究尚不完善,特别是在基准流场建立过程中忽略了三维效应、低速效应、黏性效应以及头部/前缘的钝化效应,因此其高低速气动特性均有优化设计的空间.针对此问题,本文结合高保真RANS求解器、自由变形参数化方法、鲁棒的结构网格变形方法、离散伴随方法以及序列二次规划算法,发展了基于离散伴随的宽速域飞行器气动优化设计方法.基于上述方法,针对涡波一体乘波飞行器开展了兼顾低速与高超声速气动性能的三维整机气动优化设计研究,获得了宽速域优化构型并对其进行了流动机理分析.结果表明,相较于初始构型,宽速域优化构型可以将飞行器高超声速状态下升阻特性略微提升的同时,显著增强低速状态飞行器背风面的旋涡效应,进而使飞行器低速状态的升力和升阻比均提升10%以上,改善了涡波一体宽速域乘波飞行器的高低速气动性能.     

Optimization of 3D wings based on Navier-Stokes solutions and genetic algorithms

The problem of the aerodynamic shape optimization to minimum drag, subject to geometrical and aerodynamic constraints, is considered. An accurate and computationally efficient approach to the multiobjective constrained design of 3D aerodynamic wings is proposed. The optimization is driven by full Navier-Stokes computations and Genetic Algorithms (GAs). The verification results include a variety of optimization cases for a classical test-case of ONERA M6 wing in transonic flight conditions. The method allows to significantly reduce the total drag of optimized wings, while exhibiting high robustness and keeping CFD computational volume to an acceptable level.     

Aerodynamic characteristics of airfoils with energy supply

The possibility of controlling the aerodynamic characteristics of airfoils with the help of one-sided pulsed-periodic energy supply is studied. The change in the flow structure near the airfoil and its aerodynamic characteristics are determined as functions of the magnitude of energy supply and of the energy-supply location by means of the numerical solution of two-dimensional unsteady equations of gas dynamics. It is demonstrated that external energy supply can substantially improve the aerodynamic characteristics of airfoils with a high lift-to-drag ratio. The moment characteristics of the airfoil are found.     

AERODYNAMIC OPTIMIZATION OF HIGH-SPEED TRAIN BASED ON RBF MESH DEFORMATION

基于局部型函数三维参数化方法、改进的蚁群算法和改进的克里金（Kriging）代理模型,开展了列车头型的三维气动减阻优化设计研究。为了避免复杂几何外形大变形情况下千万量级网格的重复生成,提高高速列车头型优化设计的效率,引入了缩减控制点的径向基函数网格变形技术。优化结果表明：径向基函数网格变形技术在不降低网格质量的情况下可以有效缩短网格变形的时间消耗,能够用于复杂几何外形的气动优化设计;在给定的设计空间内,控制鼻锥外形的6个关键设计参数对列车气动阻力的影响呈单调递增关系;优化后,在满足约束条件的情况下,简化外形列车的整车气动阻力减小5.41%,头尾车减阻效果明显,中间车气动阻力基本不变。     

Effect of one-sided unsteady energy supply on aerodynamic characteristics of airfoils in a transonic flow

The possibility of controlling the aerodynamic characteristics of airfoils in transonic flight regimes by means of one-sided pulsed-periodic energy supply is studied. Based on the numerical solution of two-dimensional unsteady gas-dynamic equations, the change in the flow structure in the vicinity of a symmetric airfoil at different angles of attack and the aerodynamic characteristics of the airfoil as functions of the amount of energy supplied asymmetrically (with respect to the airfoil) are determined. The results obtained are compared with the data calculated for the flow past the airfoil at different angles of attack without energy supply. It is found that a given lift force can be obtained with the use of energy supply at a much better lift-to-drag ratio of the airfoil, as compared to the case of the flow past the airfoil at an angle of attack. The moment characteristics of the airfoil are found. __________ Translated from Prikladnaya Mekhanika i Tekhnicheskaya Fizika, Vol. 49, No. 6, pp. 82–87, November–December, 2008.     

小长径比弹丸加装大展弦比尾翼的气动特性研究

为了提高小长径比弹丸射击质量,设计了一种大展弦比张开式尾翼,采用AUSM+格式、SST(shearstress transport)湍流模型和隐式算法(lower-upper symmetric Gauss-Seidel implicit method, LU-SGS),求解三维RANS 方程,对前体形状完全相同,不同展弦比的3 种尾翼弹进行了数值模拟,得到了三者在马赫数1.5~3.5 下的气动力特性的差异,分析其原因,并给出了不同展弦比张开式尾翼的适用范围. 计算结果表明:C型弹的升阻比较B 型弹在1.5 马赫数区域附近增加了7% 以上,当马赫数达到2.5 以上时,A 型弹的升阻比大于B 型弹和C 型弹,在3.5 马赫数区域附近A 型弹的升阻比较B 型弹增加了5.4% 以上. 3 种弹丸的俯仰力矩系数随着马赫数的增大而负向减少,且减少的趋势随着展弦比的增加而增大. A 型弹、B 型弹、C 型弹的静稳定裕度的变化范围分别为4%~20.3%,8.5%~23.2%,11.4%~25.6%.     

蜻蜓滑翔时柔性褶皱前翅气动特性分析

蜻蜓是自然界优秀的飞行家,滑翔是其常见且有效的飞行模式.蜻蜓优异的飞行能力来源于其翅膀的巧妙结构,褶皱是蜻蜓翅膀上最为显著的结构之一,不仅提高了翅膀的刚度,还改变了其气动特性,而飞行过程中柔性翅膀会产生变形是蜻蜓翅膀的另一特性.为揭示蜻蜓在滑翔时,柔性褶皱前翅的变形,探究褶皱和柔性的共同作用对其气动特性的影响,基于逆向工程,依据前人的测量数据和研究成果,通过三维建模软件建立了蜻蜓三维褶皱前翅的计算流体力学(computational fluiddynamics,CFD)模型和计算结构力学(computational structuralmechanics,CSD)模型,并通过模态分析验证了此模型有足够的精度.基于CFD方法和CFD/CSD双向流固耦合计算方法分别对蜻蜓滑翔飞行时刚性和柔性褶皱前翅的气动特性进行了数值模拟,结果表明,柔性褶皱前翅受气动载荷后,翅脉和翅膜产生形变,柔性前翅上下表面压力差相较于刚性前翅减小了,从而其升力和阻力也减小了,而在大攻角时,变形后的前缘脉诱导出比刚性前翅更强的前缘涡.因此在攻角小于10$^\circ$时刚性前翅的气动特性优于柔性前翅,继续增大攻角,柔性前翅的气动特性则优于刚性前翅.前翅受载后气动响应时间短,翅尖的变形最大,仅仅产生了垂直于翅膀所在平面方向上的变形,而没有发生扭转,翼根处受到应力最大,褶皱上凸部分承受蜻蜓滑翔时前翅的主要载荷.      

一种升浮一体环形浮空器气动特性研究

选取填充轻质气体的环形浮空器为研究对象,采用数值模拟方法开展高空风力发电机用浮空器气动特性研究．采用有限体积法求解不可压N-S方程和S-A湍流模型来数值模拟风力机气流场,分别对翼型,安装角,长细比,雷诺数及风力机等因素对引入浮力后浮空器气动特性影响进行研究,对比分析引入浮力后布局外形气动力特性随各外形特征参数的变化规律．数值结果表明,截面翼型弯度越大,最大升阻比越小,出现位置有一定前移;截面厚度越大,三维效应越强,最大升阻比出现有一定的滞后性;增大安装角,相当于增大攻角,使得升力系数和阻力系数随攻角变化曲线均有一定前移;引入浮力后,最大合升阻比增大,并且存在一个明显前移;长细比越小,浮空器升阻比越大,随着长细比增大,浮空器最大升阻比出现越滞后;一定范围内,雷诺数增大,浮空器动升阻比增大,引入浮力后,基于来流风速变化时,浮空器合升阻比随雷诺数增大先迅速减小然后趋于平缓,但基于浮空器尺寸变化时,合升阻比则随雷诺数增大而增大;风轮转速增大,浮空器阻力增大,升力有一定下降．     

上下反翼对双后掠乘波体低速特性的影响

相比于传统乘波体外形, 双后掠乘波体在保持高超声速良好性能的条件下能够提升乘波体低速气动性能, 但其仍存在低速稳定性不好等缺陷. 本文从密切锥乘波体理论提出给定前缘型线的乘波体设计方法, 通过给定三维前缘型线分别生成具有相同平面投影形状的上反和下反机翼双后掠乘波体. 使用CFD技术评估不同上下反程度外翼乘波体的低速性能, 分析升阻特性以及流场涡结构特点. 选取稳定性判据, 研究上下反翼对纵向和横侧向稳定性的影响. 结果表明, 机翼上下反对乘波体低速升阻特性影响较小; 不同外形均为纵向静不稳定的, 且俯仰力矩变化趋势比较类似, 机翼下反可使气动焦点位置后移, 提升纵向稳定性; 机翼上反有助于提升乘波体的横向静稳定性, 而下反则会下降; 机翼上反可以提升侧向稳定性, 且上反程度越大提升效果越明显; 同时机翼上反使乘波体的偏航动态稳定性有明显提升, 下反则会降低, 影响程度与机翼上下反程度呈正相关. 通过结果分析, 说明通过机翼上下反改善乘波体低速稳定性是可行的, 为乘波体在宽速域高超声速飞行器中的应用拓展了途径.      

On the thrust performance of a flapping two-dimensional elliptic airfoil in a forward flight

In this paper, results of numerical and experimental studies are presented for a flapping two-dimensional (2D)elliptic airfoil in a forward flight condition at a Reynolds number of 5000.The study is motivated by the experiment of Read et al. (2003), which shows that the thrust coefficient of a 2D NACA0012 airfoil deteriorated at high flapping frequency (or Strouhal number) when the induced effective angle of attack profile ceases to be a simple harmonic function in time. As to why non-simple-harmonic profile of effective angle of attack is detrimental to thrust generation is not fully understood. The paper is an attempt to address this issue by examining the flow mechanism, including near field flow structures and the associated transient aerodynamic forces and pressure field, responsible for the observed behavior. Our results show that thrust suppression can be attributed to an adverse suction effect due to high rotation rate of the airfoil and the presence of an attached leading edge vortex generated in the previous stroke. The results further show that the condition for best efficiency need not necessary coincides with the condition of best thrust performance; this observation has been made in past studies of flapping flight.     

ACTIVE SEPARATION CONTROL FOR THE FLYING-WING UAV USING SYNTHETIC JET

以小展弦比飞翼式无人机为对象,开展了基于零质量射流的主动流动控制数值模拟研究. 比较分析了应用零质量射流前后飞翼式无人机纵向气动特性的改善效果,并通过流场特征的分析探讨了流动控制技术产生气动增益的原因. 研究结果表明在模型中等迎角、大迎角范围,零质量射流技术可以显著增加升力系数,最大幅值达25%,并且拓宽了纵向力矩的线性范围. 机理分析表明,零质量射流扰动通过提高模型绕流场的边界层掺混,增强附面层内外的动量输运,使得附面层有足够的能量克服逆压梯度和黏性损耗,从而达到减缓流动分离甚至使分离流再附的目的.     

高超声速非定常流动的数值模拟与气动热计算

高超声速飞行器研究中的一个重点问题是飞行器表面的气动加热,它对飞行器的气动、热特性及安全性有重要的影响.受到当前实验技术的限制,地面实验无法准确模拟真实飞行条件,所以采用数值模拟研究气动加热问题成为目前重要的研究手段.本文采用数值方法求解三维N-S方程,得到钝头体再入模型绕流的瞬态流场,驻点温度及表面热流沿轨道变化规律.计算中采用变边界条件模拟沿轨道飞行的非定常性.     

Effects of unsteady deformation of flapping wing on its aerodynamic forces

Effects of unsteady deformation of a flapping model insect wing on its aerodynamic force production are studied by solving the Navier-Stokes equations on a dynamically deforming grid.Aerodynamic forces on the flapping wing are not much affected by considerable twist,but affected by camber deformation.The effect of combined camber and twist deformation is similar to that of camber deformation.With a deformation of 6% camber and 20°twist(typical values observed for wings of many insects),lift is increased bv 10%～20%and lift-to-drag ratio by around 10%compared with the case of a rigid flat-plate wing.As a result.the deformation can increase the maximum lift coefficient of an insect.and reduce its power requirement for flight.For example,for a hovering bumblebee with dynamically deforming wings(6?mber and 20°twist),aerodynamic power required is reduced by about 16%compared with the case of rigid wings.     

乘波体压缩面变化对其气动性能影响分析

乘波体是一种利用激波包裹特性获得高升阻比的高速飞行器构型.已有研究中,乘波体气动性能的改善主要依赖于给定源流场条件下的前缘型线优化.本文采用数值优化和计算流体力学模拟为主要手段分析了乘波体压缩面变化对其气动性能的影响,以期有效拓展乘波体的设计空间.主要内容如下:首先给出了一种基于表面局部变形的乘波体设计方法.其次结合运用增量修正参数化方法、计算流体力学分析和微分演化算法构造了乘波体压缩面外形气动优化设计流程,以一种椭圆锥形流场生成的乘波体作为基准构型开展了无黏优化.之后从优化结果中选择升阻比递增的6个典型构型进行前缘钝化处理后,基于N-S方程对其气动性能进行了评估.最后综合依据无黏/黏性计算结果分析了乘波体压缩面变化对其气动性能的影响.结果表明该部分形状的改变对乘波体气动性能影响十分明显,在升力面积不变的条件下,乘波体压缩面形状变化可导致其升阻比出现成倍变化,即使在升力不减条件下,升阻比较基准构型也可获得超过14%的提升.此外,还可导致乘波体相对压心系数出现明显偏移.