鸟类羽毛在气流中变形的力学特性研究

鸟类羽毛在飞行中的物理性质是仿生力学关心的重要问题之一.基于CFD/CSD数值模拟方法研究了羽毛微结构在气流作用下的变形和力学特征,揭示了鸟类静止时羽毛蓬松、而在飞行状态下紧贴皮肤表面保持表面光滑的物理机制.首先,通过对鸟类羽毛在显微镜下的观察,将羽毛分解成典型简单微结构以模仿羽枝单元,从而对羽毛外形和结构进行建模,之后,采用CFD/CSD方法分析比较了两种典型羽枝模型结构(片状和枝状羽枝单元)的变形和力学特征,最后,基于上述片状羽枝模型进一步研究了来流方向对羽枝变形的影响机理及多根排列羽枝的变形和力学特征.结果表明:在一定风向的范围内,羽毛在气流下都具有保持紧贴皮肤表面的变形趋势,这种紧贴壁面的趋势只有在气流与羽轴几乎垂直时才会改变;在来流侧滑角为45°时,羽枝沿皮肤表面法向下压的变形最为显著,尖端位移达原始高度的约97%;多根排列的羽枝在顺流方向气动载荷逐渐下降,与迎风首根羽枝最大差距约11%.此研究工作对于理解鸟类飞行时羽毛的力学特性有明确的学术价值.

转捩位置对全动舵面热气动弹性的影响

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=15 km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.

展向动量测定法与前掠翼流动机理研究

为揭示前掠翼与后掠翼的流动差异,研究前掠翼流动的特点和机理,设计了可进行直接比较的具有相同翼型剖面、相同展弦比、无根梢比的前掠45°(Λ=-45°)与后掠45°(Λ=45°)机翼模型,采用基于雷诺平均N-S方程流场求解器对前掠和后掠翼低速纵向气动性能进行数值模拟计算,并提出了展向动量测定法前掠翼流动机理进行了深入分析,研究结果表明:1)前掠翼展向动量输运使得升力向翼根汇聚,前掠翼展向升力分布更接近于椭圆分布,致使前掠翼诱导阻力更小;2)黏性对前掠翼的流场计算影响很大,文献中基于势流理论得到前掠翼比后掠翼气动性能好的结果是不准确的;3)在小迎角时,前掠翼气动效率与后掠翼相当,仅在最大升阻比迎角时前掠翼优于后掠翼,中等迎角下前掠翼翼根分离导致气动效率下降,但前掠翼具有更好的大迎角失速特性,有利于前掠翼大迎角飞行;4)相同总升力的情况下,前掠翼的翼根弯矩只有后掠翼的翼根弯矩的89.4%,采用前掠翼更有利于减轻机翼的结构重量。

高超声速飞行器气动弹性的近期进展与发展展望

近年来高超声速飞行器气动弹性(尤其是热气动弹性)问题的关注度不断增强,相关内容已成为学术界的重点研究方向。本文旨在总结高超声速气动弹性问题的近期研究成果,并对进一步发展做出展望。首先从三个方面综述了气动弹性的计算方法:气动力的计算、常温下气动弹性求解方法及热环境下气动弹性的研究方法。然后,按照工程实际中出现的相关问题,将高速飞行器的热气动弹性问题分为与外流相关的飞行器热气动弹性问题以及与内流道相关的热气动弹性问题,分别探讨其研究情况及未来发展。其中,比较详细地论述了围绕发动机相关的热气动弹性问题(包括冲压发动机内部的壁板颤振和火箭发动机喷管的气动弹性),指出该问题有可能成为未来一个重要的研究范畴。

柔性仿羽毛结构抑制边界层转捩的初步探索

通过观察,发现存在展向流动时鸟类柔性的羽毛侧缘会卷起,这对展向流动起到了额外的阻碍作用,即流向和展向的阻力特征是不同的。为了研究羽毛表面的这种各向异性阻力特征对边界层转捩流动的影响,本文从宏观角度出发建立了一种唯象力学模型来刻画表面的各向异性。然后,运用直接数值模拟的方法研究了该模型对平板边界层转捩的影响。自由来流马赫数为0.2,基于入口处边界层排移厚度的雷诺数为732。研究结果表明,在模型抑制展向流动的作用下,平板边界层转捩明显推迟,不同的参数下至少可以推迟一倍以上距离。即使在流动进入湍流状态后,壁面的摩擦阻力系数也大大降低,维持在与层流摩擦系数相当的水平。转捩位置的推迟和湍流区摩擦阻力的降低都有利于降低平板的阻力。如果考虑卷起的羽毛侧缘对流向流动带来额外阻力,这会使得转捩提前,减阻的效果降低。此研究结果一方面揭示了鸟类飞行的部分奥秘,另一方面也为边界层被动控制措施提供了一种新的思路。

基于微米力学实验的页岩Ⅰ型断裂韧度表征

页岩断裂韧度KIC是页岩气储层水力压裂设计的基础参数之一,由于组成的非均质性,常规宏观力学测量方法存在制样困难、力学解释参数不连续、精度偏低等问题.如何及时获取页岩的断裂特性,确保安全高效的工程施工,是当前面临的一大问题.因此,提出了基于微米力学实验的页岩Ⅰ型断裂韧度分析方法,可用于页岩微裂纹起裂、发育直至形成宏观裂纹的机理研究,进行页岩宏观Ⅰ型断裂韧度预测.基于页岩多尺度组成分析,开展了维氏压头和玻氏压头的页岩微米力学实验,分析了页岩残余压痕与压头间的相似关系、有效测试载荷以及压头参数的优化与选择.分析了不同压入载荷下的页岩细观断裂韧度分布特征,开展了宏观巴西圆盘实验,验证页岩微米力学测试方法的适用性.研究结果表明,在有效载荷范围内的页岩细观Ⅰ型断裂韧度波动性较小,当压入载荷过大时,由于岩样压痕区域出现局部剥落导致断裂韧度测量值偏小.与宏观实验的比对分析显示,微米力学实验的KIC平均值为0.86 MPa·√m,直槽切缝巴西圆盘实验得到的KIC平均值为0.92 MPa·√m,两类方法的统计平均值较为接近,页岩局部组成的非均质性使得微米力学测量结果较宏观测试更为分散.研究结果可用于页岩宏观Ⅰ型断裂韧度预测,为有效解决页岩气储层水力压裂参数评价提供新的思路和方法.

展向自适应机翼总体气动特性分析

针对展向自适应机翼的气动特性随折叠角度变化的问题,以经典翼型NACA0012为基础,设计了内外段比例为7∶1的展向自适应机翼。基于结构化网格和雷诺平均N-S方程,采用自主开发的流场求解器,研究了自适应机翼在不同速域、不同折叠角度情况下的总体气动性能以及操纵特性。从升阻比和机翼表面压力分布两个方面,对比了外段机翼在不同折叠角度下的总体气动效率以及折叠角度对流场特性的影响规律。研究结果表明,自适应机翼的对称变形在合适的折叠角度下可以使亚声速和超声速飞行条件下的气动效率大幅增加,增幅高达28%;亚声速飞行时的高气动效率来源于升力增加和阻力减小的共同作用,而超声速时的高气动效率主要来源于阻力的减小;在跨声速飞行条件下的气动特性随折叠角度变化不明显;非对称变形可以产生明显的用于方向操纵的滚转力矩和偏航力矩。通过将外段机翼折叠到不同角度,展向自适应机翼可以适应不同的飞行工况,获得更好的气动效益,可应用于下一代亚声速或超声速飞机。

某战术导弹舵面异常高频抖动机理分析

通过CFD数值模拟分析某战术导弹舵面高频抖动机理。该导弹采用正常式气动布局,舵面靠近弹尾,发动机喷管位于弹体底部中央。在亚音速飞行试验中,固体发动机工作结束时,舵面出现异常高频抖动现象,机理不明。数值模拟分析结果表明:发动机工作结束时,弹体底部非定常分离流动随着发动机尾喷流减弱而增强,舵面在弹体底部非定常气动力激励下高频抖动。

降低超声速飞行器声爆的一种主动控制方法

声爆抑制问题是新一代环保型超声速客机研发过程中亟需解决的关键技术瓶颈之一。提出了一种降低超声速飞行器声爆的主动控制方法,主要是在翼型下表面前缘附近处开口进行吸气,同时在翼型下表面后缘附近进行喷气,且保持吸气量与喷气量相等。分别以菱形翼型和NACA0008翼型作为基准算例,研究了吸气口与喷气口位置、气体质量流量以及飞行迎角对地面声爆最大过压、阻力系数以及升阻比的影响。结果表明该方法能够显著地降低超声速飞行器声爆水平以及阻力,并且后缘处喷气比前缘处吸气更容易达到降噪的效果。将主动控制方法应用于基准菱形(NACA0008)翼型后,当气流质量流量为6.5 kg/s(7.5 kg/s)时,最大正过压值降低了12.87%(12.85%),最大负过压值的绝对值降低了33.83%(56.77%),阻力系数降低了9.50%(10.96%)。可见文中提出的低声爆方法效果良好,为新一代低声爆超声速客机的设计提供了有益的参考。

振荡激波作用下受热壁板主共振特性分析

冲压发动机内部的激波串往往存在振荡特征,研究振荡激波作用下受热壁板的主共振特性对结构安全性具有重要指导作用。基于Von-Karman大变形理论和当地活塞流理论,采用Galerkin方法建立了振荡激波作用下壁板振动的动力学模型,通过龙格-库塔法对非线性动力学方程进行数值计算获得系统非线性振动响应,发现受热壁板在振荡激波作用下存在振动突跳和双稳态现象等典型的非线性动力学行为,并分析了振荡激波强度、激波振荡幅值、温度、激波振荡的中心位置、来流马赫数对系统振动突跳和双稳态特性的影响。结果表明:随着振荡激波强度、激波振荡幅值和来流马赫数增大,系统共振峰值不断单调增大,而双稳态区间由不存在先变宽、再变窄直至消失,然后再急剧变宽;温度增大会对系统产生“刚度弱化效应”;将激波振荡的中心位置向壁板两端移动可有效地降低系统共振峰值并抑制振动突跳和双稳态现象。

高超声速飞行器非线性纵向运动稳定性研究

针对研究高超声速飞行器飞行动力学稳定性时忽略方程非线性影响的问题,建立了保留非线性特征的六自由度动力学方程组和非线性活塞理论耦合模型。研究了初始迎角和活塞理论非线性项对高超声速飞行器纵向稳定性的影响,并在稳定巡航中施加瞬时扰动,分析其姿态响应。仿真结果表明:高超声速飞行器姿态收敛速度比常规低速飞行器慢很多;对迎角施加扰动后,飞行器姿态变化主要影响短周期模态,且偏离平衡位置越远,收敛越快;改变气动力非线性项时,其姿态变化主要影响长周期模态,且收敛较慢;两种扰动均对振荡周期无明显影响。

Investigation on the mechanism of aeroelastic hazard during ground test of rocket nozzle

Side loads and aeroelastic stability of rocket nozzle were studied by solving Navier-Stokes equation coupled with structural equation of motion.The computation was implemented at different total pressure inlet conditions,and flow phenomena of free shock separation(FSS) and restricted shock separation(RSS) were captured.At certain total pressure inlet conditions,it was found that both kinds of separations existed in nozzle flow filed,while RSS exhibited combined space asymmetry and time unsteady characteristics.The corresponding asymmetric circumferential pressure distribution,strong pressure fluctuation in separation region and large range of displacement of shock wave all led to severe side loads.Besides,for flexible nozzles,the low pressure gradient in separation region might reduce structure stability at nozzle exit,resulting in large local deformation.It was also found that aeroelasticity exhibited buffeting characteristic due to the asymmetric separation,resulting in reduction of aeroelastic stability,even structure destruction.Moreover,aeroelasticity might amplify side loads and aggravate its growth rate.However,with increment of inlet pressure,nozzle aeroelastic stability was also increased when a full flow was nearly reached.

An Improved LU-SGS Implicit Scheme for High Reynolds Number Flow Computations on Hybrid Unstructured Mesh

The lower-upper symmetric Gauss-Seidel （LU-SGS） implicit relaxation has been widely used because it has the merits of less dependency on grid topology, low numerical complexity and modest memory requirements. In original LU-SGS scheme, the implicit system matrix is constructed based on the splitting of convective flux Jacobian according to its spectral radius. Although this treatment has the merit of reducing computational complexity and helps to ensure the diagonally dominant property of the implicit system matrix, it can also cause serious distortions on the implicit system matrix because too many approximations are introduced by this splitting method if the contravariant velocity is small or close to sonic speed. To overcome this shortcoming, an improved LU-SGS scheme with a hybrid construction method for the implicit system matrix is developed in this paper. The hybrid way is that： on the cell faces having small contravariant velocity or transonic contravariant velocity, the accurate derivative of the convective flux term is used to construct more accurate implicit system matrix, while the original Jacobian splitting method is adopted on the other cell faces to reduce computational complexity and ensure the diagonally dominant property of the implicit system matrix. To investigate the convergence performance of the improved LU-SGS scheme, 2D and 3D turbulent flows around the NACA0012 airfoil, RAE2822 airfoil and LANN wing are simulated on hybrid unstructured meshes. The nu- merical results show that the improved LU-SGS scheme is significantly more efficient than the original LU-SGS scheme.

Reliability and Sensitivity Analysis of Transonic Flutter Using Improved Line Sampling Technique

The improved line sampling （LS） technique, an effective numerical simulation method, is employed to analyze the probabilistic characteristics and reliability sensitivity of flutter with random structural parameter in transonic flow. The improved LS technique is a novel methodology for reliability and sensitivity analysis of high dimensionality and low probability problem with implicit limit state function, and it does not require any approximating surrogate of the implicit limit state equation. The improved LS is used to estimate the flutter reliability and the sensitivity of a two-dimensional wing, in which some structural properties, such as frequency, parameters of gravity center and mass ratio, are considered as random variables. Computational fluid dynamics （CFD） based unsteady aerodynamic reduced order model （ROM） method is used to construct the aerodynamic state equations. Coupling structural state equations with aerodynamic state equations, the safety margin of flutter is founded by using the critical velocity of flutter. The results show that the improved LS technique can effectively decrease the computational cost in the random uncertainty analysis of flutter. The reliability sensitivity, defined by the partial derivative of the failure probability with respect to the distribution parameter of random variable, can help to identify the important parameters and guide the structural optimization design.

Experimental and computational investigation of novel vertical tail buffet suppression method for high sweep delta wing

Our research aim is to investigate the buffet alleviation effect of static or vibrating bulges attached on the forebody surface of the model.Experiments and numerical simulations on a model consisting of a sharp-edged,70°-leading edge sweep delta wing and twin swept back vertical tails were conducted.Models with different bulges were tested and computed at 10 and 20 m/s of free stream velocity at angles of attack ranging from 20°–50°.Dynamic strain gauge and multichannel data acquisition and analysis system were employed for the measurement of unsteady root strain on the vertical tails.Experimental and computational results show that both static and vibrating bulges behave effectively as a novel tool to alleviate tail buffet,and the alleviation effect depends largely on the vibrating frequency.Besides,one-sided bulge can only alleviate the buffeting response for the tail of the same side,and it has no obvious alleviation effect for the opposite tail.Results of the spectral analysis reveal that there are generally three peaks of spectral density for aircrafts of this configuration,and bulges used in this paper could alleviate tail buffeting,but the total lift and drag of the whole model show no obvious deviation compared to the base model and the dominant frequency of the vibration of the tails has not shifted.

Numerical study on the correlation of transonic single-degree-of-freedom flutter and buffet

Transonic single-degree-of-freedom(SDOF) flutter and transonic buffet are the typical and complex aeroelastic phenomena in the transonic flow. In this study, transonic aeroelastic issues of an elastic airfoil are investigated using Unsteady Reynolds-Averaged Navier-Stokes(URANS) equations. The airfoil is free to vibrate in SDOF of pitching. It is found that, the coupling system may be unstable and SDOF self-excited pitching oscillations occur in pre-buffet flow condition, where the free-stream angle of attack(AOA) is lower than the buffet onset of a stationary airfoil. In the theory of classical aeroelasticity, this unstable phenomenon is defined as flutter. However, this transonic SDOF flutter is closely related to transonic buffet(unstable aerodynamic models) due to the following reasons. Firstly, the SDOF flutter occurs only when the free-stream AOA of the spring suspended airfoil is slightly lower than that of buffet onset, and the ratio of the structural characteristic frequency to the buffet frequency is within a limited range. Secondly, the response characteristics show a high correlation between the SDOF flutter and buffet. A similar "lock-in" phenomenon exists, when the coupling frequency follows the structural characteristic frequency. Finally, there is no sudden change of the response characteristics in the vicinity of buffet onset, that is, the curve of response amplitude with the free-stream AOA is nearly smooth. Therefore, transonic SDOF flutter is often interwoven with transonic buffet and shows some complex characteristics of response, which is different from the traditional flutter.

The CDF and its sensitivity analysis of stochastic structure with stochastic excitation by advanced stratified line sampling

For the stochastic structure with stochastic excitation, an advanced stratified line sampling (SLS) method is presented to obtain the cumulative distribution function (CDF) of the structural response and its sensitivity. The advanced stratified line sampling method introduces a set of middle failure subsets firstly. And for each subset, the conventional line sampling can be used to obtain the corresponding value of the response's CDF. At the same time, the sensitivity estimations of each failure subset can also be computed by modifying the important direction and corresponding reliability coefficients. The properties of CDF sensitivity are proved while the performance function is linear with normal random variables. After two simple examples are used to demonstrate the properties of CDF sensitivity and the feasibility of the presented method, the method employed to analyze the CDF and corresponding sensitivity of root bending moment (RBM) responses for the stochastic BAH is wing with gust excitation to a square-edged gust and to a Dryden gust. The results show that the parameters of the second and the fifth order modals exert more influence on the CDF of response than the other ones, and the presented SLS method can more significantly reduce the computational cost compared with Monte Carlo simulation (MCS).

Numerical simulation of rotor-airframe aerodynamic interaction based on unstructured dynamic overset grids

A method of unstructured dynamic overset grids is developed for the numerical simulation of helicopter unsteady rotorairframe aerodynamic interaction.For the effective treatment of the relative motion between the rotor and the airframe,the domain of flowfield is divided into two overset subzones,namely,a rotational subzone containing the blades and a stationary subzone containing the airframe.The overset part of two subzones is used to convect the flow variables of the two zones.The Taylor series expansion is used to obtain a second-order spatial accuracy,and dual-time stepping is adopted to improve the solution accuracy.Mesh deformation from the blade motion in forward flight is treated by using a spring analogy.Validation is made by numerically simulating the flows around a wind tunnel configuration and comparing the predicted time-averaged and instantaneous inflow and airframe surface pressure distributions with the experimental data.It shows that the present method is efficient and robust for the prediction of complicated unsteady rotor-airframe aerodynamic interaction phenomena.

Characteristic analysis of lock-in for an elastically suspended airfoil in transonic buffet flow

Numerical simulations are performed to study the aeroelastic responses of an elastically suspended airfoil in transonic buffet flow, by coupling the unsteady Reynolds-averaged Navier- Stokes （RANS） equations and structural motion equation. The current work focuses on the char- acteristic analysis of the lock-in phenomenon. Great attentions are paid to studying the frequency range of lock-in and the effects of the three parameters, namely the structural natural frequency, mass ratio and structural damping, on lock-in characteristic of the elastic system in detail. It is found that when the structural natural frequency is close to the buffet frequency, the coupling fre- quency of the elastic system is no longer equal to the buffet frequency, but keeps the same value as the structural natural frequency. The frequency lock-in occurs and stays present until the structural nature frequency is near the double buffet frequency. It means that the lock-in presents within a broad range, of which the lower threshold is near the buffet frequency, while the upper threshold is near the double buffet frequency. Moreover, the frequency range of lock-in is affected by mass ratio and structural damping. The lower the mass ratio and structural damping are, the wider the range of lock-in will be. The upper threshold of lock-in grows with the mass ratio and structural damping decreasing, but the lower threshold always keeps the same.

Design of a new tandem wings hybrid airship

It is scientifically important science value and engineering promising to develop the buoyancy-lift integrated hybrid airship for high attitude platform.Through the numerical method,a new tandem wings hybrid airship with both higher utility value and economy efficiency was obtained and its total performance and technical parameters were analyzed in detail.In order to further improve the lift-drag characteristics,we implemented the optimization design for aerodynamic configuration of tandem wings hybrid airship via the response surface method.The results indicate that the tandem wings hybrid airship has considerable volume efficiency and higher aerodynamic characteristics.After optimization,the lift-drag ratio of this hybrid airship was increased by 6.08%.In a given gross lift condition,tandem wings hybrid airship may provide more payload and specific productivity.Furthermore,the size of tandem airship is smaller so the demand for skin flexible materials can be reduced.Results of this study could serve as a new approach to designing buoyancy-lifting integrated hybrid airship.