Aerodynamic characteristics for a passenger aircraft model with harmonic oscillations on rolling and yawing angles at high angles of attack

Experimental values of complexes of aerodynamic derivatives were obtained in a wind tunnel and analyzed for subsonic flow of a passenger aircraft model with harmonic oscillations on the rolling and yawing angles. It was demonstrated that for near-critical angles of attack, the complexes of aerodynamic derivatives taken for rolling and yawing have a string dependency on the normalized oscillation frequency. It was proven that this dependency is driven by derivatives with respect to angular velocities and angular acceleration. A mathematical model for aerodynamic loads was developed; it can be applied for aircraft flight dynamic problems and has satisfactory approximation to experimental aerodynamic characteristics.     

基于AC-DBD等离子体激励的NACA0012翼型动态失速控制

以NACA0012翼型为研究对象,采用动态测压及PIV测量技术,研究了AC-DBD等离子体激励器对翼型俯仰及耦合运动动态失速的控制作用和机理.研究表明,等离子体激励能够显著推迟失速迎角,抑制失速后的升力系数陡降,提前流动再附和升力系数回升,减小升力及俯仰力矩系数曲线迟滞环面积,改善翼型气动特性.研究了不同运动参数及激励...     

高升阻比模型天平校准安装角误差控制要求

文章针对高升阻比类模型气动力测量的要求,对天平校准安装角误差影响进行了初步分析,天平的单元校准中,主要有两个安装角对天平校准结果产生影响,这两个安装角分别为攻角和滚转角,研究结果表明:天平的不同分量和不同校准系数对安装角误差的要求是不尽相同的,力分量以及滚转力矩校准,安装角误差可以允许在度水平上;俯仰力矩和偏航力矩校准...     

飞行器大攻角非定常气动特性神经网络建模

大攻角气动特性预测与气动建模是新型飞行器提升飞行性能的重要内容.以轴对称导弹简化模型为研究对象，首先采用计算流体力学方法，对70°大攻角状态的非定常气动特性进行数值模拟，计算方法基于RANS的N-S方程，湍流模型采用SA模型，对流场采用有限体积法离散，无黏项采用Roe通量差分分裂格式，黏性项采用中心差分，时间推进采用LU-SGS格式的双时间步法.飞行器运动模式采用强迫振荡的方式，对5种不同振荡频率进行了非定常数值计算，并记录每一内迭代周期最终的气动力和力矩数值.其次，以CFD预测结果作为气动建模的样本，采用动导数模型、多项式模型等传统方法，进行气动建模，并分析其有效性和精度.最后采用神经网络方法对大攻角非定常气动力进行建模，并和动导数模型、多项式模型进行精度对比.结果表明，基于神经网络的人工智能气动建模方法具有较高的精度和适应性.该方法为飞行器大攻角非定常非线性气动建模，大攻角飞行稳定性分析与控制提供理论参考.      

Influence of small angles of attack and sweep on the rolling moment at a hypersonic flow around the bodies of revolution

The influence of small attack and sweep angles on the rolling moment variation at a supersonic flow around the bodies close to the bodies of revolution is considered. Based on the method of the differential hypothesis of locality the integral expressions are obtained for the derivatives of the rolling moment coefficient with respect to the attack and sweep angles. The norm of these derivatives is defined and the estimates for the norm are obtained.     

钝锥再入飞行器黏性干扰效应建模

针对典型的钝锥外形, 采用统一气体动理学格式(UGKS)模拟了高度70~110 km下不同Mach数和攻角的流场, 进行了流场特性的分析, 并基于黏性干扰的理论成果, 将气动力特性与第3黏性干扰参数、攻角和Mach数等参数进行关联, 建立了气动力系数的黏性干扰模型, 给出了模型预测结果的相关性分析和准确性评估。经初步测试, 该模型预测结果与UGKS直接模拟结果具有良好的一致性, 对工程应用快速获取高空气动特性具有重要意义。      

路堤倾角变化对高速列车侧风运行气动特性影响

采用计算流体力学方法,探讨路堤倾角变化对高速列车侧风运行气动特性影响.结果表明:随着路堤倾角增大,头车侧向力系数先增加后减小,而中间车侧向力系数则持续减小.尾车上的侧向力系数先是略微减小,当路堤倾角达到47.5°时,侧向力系数随倾角增加而增加.作用在中间车上的负升力随着路堤倾角增加而增加.尾车的升力系数对路堤倾角变化非常敏感,变化幅度较大.路堤倾角变化时高速列车绕流流场变化,从而导致列车受到的气动力发生变化.     

Aerodynamic derivatives of the model of the jettisonable module of the emergency rescue system at hypersonic velocities

The model, experimental equipment, and test program are briefly described. A method of determining the aerodynamic characteristics of the model on the facility with free oscillations is presented. Aerodynamic derivatives of the pitching moment of the model are obtained for two positions of the axis of rotation and Mach numbers M_∞ = 2, 4, and 6. At M_∞ = 2, the model with the rear position of the axis of rotation is not balanced at low angles of attack, whereas irregular self-sustained oscillations of the model with the frontal position of the axis of rotation arise.     

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The influence of the coil pitch on the laser oscillation parameters of copper ion lines using a helical hollow cathode design are investigated. A rapid growth of the laser gain with increasing pitch is demonstrated. A 40% per meter gain coefficient of the 780.8 nm copper ion line at 20 A peak current is measured. Simultaneous oscillation on 7 IR copper ion lines and 2 krypton ion lines from 15 cm active length is obtained.     

振荡射流改善翼型气动性能的实验研究

本文针对采用振荡射流控制流动分离改善大攻角下翼型气动特性的问题,在NACA0015翼型上进行了多种工况的风洞实验。结果表明:在失速攻角附近,振荡射流抑制流动分离提高翼型升力系数的作用十分明显,可将翼型失速攻角推迟2°左右。存在最佳的振荡射流频率段、射流动量范围和射流位置,使得翼型性能的改善最大。实验还得到了振荡射流的频率、动量和施加位置等参数对翼型气动性能的影响规律。     

How bumps on whale flippers delay stall: an aerodynamic model

Wind tunnel experiments have shown that bumps on the leading edge of model humpback whale flippers cause them to "stall" (i.e., lose lift dramatically) more gradually and at a higher angle of attack. Here we develop an aerodynamic model which explains the observed increase in stall angle. The model predicts that as the amplitude of the bumps is increased, the lift curve flattens out, leading to potentially desirable control properties. We find that stall delay is insensitive to the wavelength of the bumps, in accordance with experimental observations.     

RELATIONSHIP BETWEEN INTENSITY OF SURFACE SECOND HARMONIC GENERATION AND MOLECULAR STRUCTURE IN HELIX MOLECULAR FILMS

The relationship between intensity of surface Second Harmonic Generation (SHG) and the pitch, radius and azimuthal angle of molecules on the surface of helix molecular films is analyzed. The values of pitch and radius corresponding to the maximal intensity of SHG have been shown to exist. The theoretical results are compared with the experimental ones reported by M?hlman et al., and the limitation of their conclusion is also pointed out.     

RELATIONSHIP BETWEEN INTENSITY OF SURFACE SECOND HARMONIC GENERATION ANDMOLECULAR STRUCTURE INHELIX MOLECULAR FILMS

The relationship between intensity of surface Second Harmonic Generation (SHG) and the pitch, radius and azimuthal angle of molecules on the surface of helix molecular films is analyzed. The values of pitch and radius corresponding to the maximal intensity of SHG have been shown to exist. The theoretical results are compared with the experimental ones reported by Mohlman et al., and the limitation of their conclusion is also pointed out.     

Effect of relative camber on the aerodynamic performance improvement of asymmetrical blunt trailing-edge modification

In this paper, the aerodynamic performance of the S series of wind turbine airfoils with different relative cambers and their modifications is numerically studied to facilitate a greater understanding of the effects of relative camber on the aerodynamic performance improvement of asymmetrical blunt trailing-edge modification. The mathematical expression of the blunt trailing-edge modification profile is established using the cubic spline function, and S812, S816 and S830 airfoils are modified to be asymmetrical blunt trailing-edge airfoils with different thicknesses. The prediction capabilities of two turbulence models, the k-ω SST model and the S-A model, are assessed. It is observed that the k-ω SST model predicts the lift and drag coefficients of S812 airfoil more accurately through comparison with experimental data. The best trailing-edge thickness and thickness distribution ratio are obtained by comparing the aerodynamic performance of the modifications with different trailing-edge thicknesses and distribution ratios. It is, furthermore, investigated that the aerodynamic performance of original airfoils and their modifications with the best thickness of 2% c and distribution ratio being 0:4 so as to analyze the increments of lift and drag coefficients and lift–drag ratio. Results indicate that with the increase of relative camber, there are relatively small differences in the lift coefficient increments of airfoils whose relative cambers are less than 1.81%, and the lift coefficient increment of airfoil with the relative camber more than 1.81% obviously decreases for the angle of attack less than 6.3°. The drag coefficient increment of S830 airfoil is higher than that of S816 airfoil, and those of these two airfoils mainly decrease with the angle of attack. The average lift–drag ratio increment of S816 airfoil with the relative camber of 1.81% at different angles of attack ranging from 0.1° to 20.2° is the largest, closely followed by S812 airfoil. The lift–drag ratio increment of S830 airfoil is negative as the angle of attack exceeds 0.1°. Thus, the airfoil with medium camber is more suited to the asymmetrical blunt trailing-edge modification.     

光学炮塔气动载荷的非稳态特性

采用耦合J-B模型的IDDES模型与双时间步LU-SGS方法开展炮塔非稳态气动载荷的数值仿真研究。炮塔流动会发生分离,天顶位置的分离角大于90°;当流动绕过炮塔时,形成马蹄涡、脱落涡街等非稳态流场结构,导致气动载荷也具有非稳态特性;炮塔顶点的脉动静压功率谱在1.6~40.0 kHz进入各向同性均匀湍流的惯性子区,基本满足Kolmogrov的-5/3定律;气动力以阻力为主,横向力的脉动幅值大,气动力矩则以俯仰力矩为主,滚转力矩的脉动幅值大,偏航力矩可以忽略不计;气动力和力矩的功率谱主要集中在1 kHz以下,存在多个尖峰频率,主频约为230 Hz (斯特劳哈尔数为0.15)。在ATP系统设计之初,需要考虑光学炮塔所受气动载荷的非稳态特性,并规避尖峰频率尤其是主频的谐振破坏问题。     

进入轨道偏差对火星科学实验室气动力特性的影响

针对火星科学实验室(MSL)高超声速进入过程,利用三维并行程序求解流体力学Navier-Stokes方程,耦合真实气体模型,分析火星大气中真实气体效应对进入器气动力特性的影响量在进入轨道发生偏差时的变化规律.结果表明:对海盗号的计算结果与飞行数据符合很好,验证了火星大气真实气体模型和计算方法;真实气体效应影响下,激波层厚度大为减小,温度下降明显,进入器阻力系数明显增加,升力系数变化不大,俯仰力矩系数增加,基准状态下配平攻角较完全气体减小约2.2°;高度不变,Ma数增加导致阻力系数和俯仰力矩系数增大,配平攻角和完全气体的差值由1.6°增加到2.6°,表明Ma数变大时真实气体效应引起的气动力变化增强;Ma数不变,高度增加略微减弱波后化学反应,对进入器气动力特性基本没有影响.     

Investigation of the aerodynamic properties and flow field around hypervelocity objects in a ballistic test range

The equipment of a ballistic test range for research on hypersonic motion and flow around typical models of explosion-propelled projectiles. A method for obtaining instantaneous interferograms of the flow around an object and a procedure for calculating the radial density distributions from measurements made on them are briefly set forth. The values of the aerodynamic drag at zero angle of attack are obtained for three models in trajectory experiments, and a comparative assessment of their stability is made. The principles for performing numerical aerodynamic calculations of hypervelocity flow around objects typically having a compound shape are formulated. Zh. Tekh. Fiz. 69, 6–11 (December 1999)     

交叉旋翼悬停气动性能和流场干扰

针对交叉旋翼复杂的气动干扰问题,建立了一种适合于交叉旋翼气动分析的数值模拟方法.该方法采用三维非定常Reynolds平均Navier?Stokes(RANS)方程来求解流场,使用动态嵌套网格方法模拟旋翼运动.使用共轴旋翼悬停实验结果验证了该方法的准确性.利用该方法模拟了交叉旋翼在不同状态下的流场,计算了拉力和悬停效率,...     

直升机旋翼桨叶气动力学参数飞行测试系统

根据某型号直升机定型试飞的需要,研制出了动态测量旋翼桨叶气动力学参数的光电测试系统。其中运动参数(挥舞角、摆振角和变距角)由二维位置敏感传感器响应桨叶表贴发光点的位置变化测得,挥舞应变由光纤光栅传感器测得。利用自主搭建的桨叶运动姿态仿真平台开展了相关测试实验。结果表明:各角度值与对应输出电压间有很好的线性关系,测量范围内的角度测量误差小于1%;光纤光栅与电阻应变片对比测试的一致性较好,线性相关系数达到了0.999;光纤光栅应变测量的复现性很好,两次应变测量的相对偏差仅为1.22%。系统具有较高的测试精度。     

串列扑翼间前缘涡在不同间距下对平均推力的影响

利用压力传感器测量扑翼的瞬时力,利用数字粒子测速仪（digital particle image velocimetry,DPIV）系统测量扑翼的前缘涡以及周围的流场,来揭示前缘涡在不同间距下对扑翼平均推力的影响.实验在一个低Reynolds数循环水洞中进行,两串列扑翼均做二维正弦平动.在固定的相位差下,当间距增加时,后翅前缘涡对前翅的影响具有相似性,均提高或者均降低前翅的平均推力.前翅平均推力的提高是由于后翅的前缘涡提高了前翅尾部的射流速度以及有效攻角.随着间距的增加,后翅前缘涡对前翅的影响急剧下降,使得前翅的平均推力快速接近于单翼值.在固定的相位差下,当间距增加时,前翅的脱落涡对后翅的影响变化非常大,后翅的平均推力可能先升高后降低,这是因为间距改变了前翅脱落涡作用于后翅的时间点.当前翅脱落涡遇到后翅,并且和后翅的前缘涡有相同的旋转方向时,前翅的脱落涡会抑制后翅前缘涡的形成,并且后翅的有效攻角减小,其平均推力降低.如果这两个涡的旋转方向相反,那么后翅有效攻角就会增大,平均推力值就会提高.