一种前后缘连续变弯度翼型气动特性分析

提出了一种基于厚度不变的翼型前后缘连续偏转变形规律,并以NACA0015翼型为例,实现了翼型变弯度的参数化。以柔性伸缩蒙皮支撑结构和机械结构实现了可连续光滑偏转后缘的变弯度翼型,验证了变形规律。以前后缘偏转角度为参数,计算并分析了各个变形状态下翼型扰流流场和气动特性,讨论了变形参数对气动特性的影响,研究了气动特性的产生机理。研究结果表明:在大迎角下,前缘偏转角对翼型失速有一定的抑制作用;在中小迎角范围内,翼型升阻比随着后缘偏转角的增大而增大,且不论迎角如何变化,总可以通过前后缘偏转来获得较高的升阻比。     

可连续光滑偏转后缘的变弯度翼型气动特性分析

以变弯度翼型为研究对象,计算了其六种外形的绕流流场,分析了不同的连续光滑变形翼型与传统偏转翼型的气动特性,讨论了变形参数对气动特性的影响,研究了气动特性的产生机理;与此同时,以形状记忆聚合物柔性蒙皮和机械结构实现了可连续光滑偏转后缘的变弯度翼型,并在风洞实验中测试了其气动特性。计算和实验结果表明:可连续光滑偏转后缘的变弯度翼型能改进传统主翼-简单襟翼翼型的气动特性和流场分离特性;可变形段范围、转轴位置、后缘偏转角度、后缘高度等变形参数对气动特性具有显著影响。     

GAW-1翼型前后缘变弯度气动性能研究

传统增升装置主要用于提高飞机起降气动性能。利用计算流体力学(CFD)的方法,引入了通用飞机翼型的前后缘变弯装置的概念,数值模拟了GAW-1翼型在爬升状态时,前缘变弯装置、后缘襟翼/副翼偏转以及前后缘装置综合偏转对翼型气动特性的影响。研究表明,前缘变弯装置可以有效地改善翼型的失速特性,失速迎角提高了3°左右,最大升力系数提高了4.56%;同时提高升阻比50%~120%;但在设计升力系数下,升力系数和阻力系数都略微减小。另一方面,后缘变弯装置可以改变最大升阻比所对应的迎角,以及在小迎角时,提高升力系数6%左右。翼型综合偏转可以在小迎角时增加升力系数,在大迎角时增加升阻比。     

后缘连续变弯度对跨声速翼型气动特性的影响

针对后缘连续变弯度对跨声速翼型气动特性的影响进行了研究。首先不考虑翼型后缘连续变弯度,基于搭建的优化设计系统对跨声速翼型进行气动减阻优化设计,通过添加不同的约束优化得到两种跨声速翼型:无激波翼型和超临界翼型。然后在这两种翼型的基础上,以后缘偏转角度为设计变量、以阻力系数最小为目标,针对不同的升力系数分别进行优化设计,并根据优化结果深入分析后缘连续变弯度对这两种翼型极曲线特性的影响机理。优化结果表明:无激波翼型与超临界翼型相比,其设计点处的气动特性较好,但鲁棒性较差;升力系数小于设计升力系数时,应用后缘连续变弯度后,无激波翼型的极曲线特性明显提高,减阻最高达到3.9%,而超临界翼型的极曲线特性提高不明显;升力系数大于设计升力系数时,应用后缘连续变弯度后,无激波翼型和超临界翼型的极曲线特性都明显提高,减阻分别达到2.4%~18.1%和1.7%~13.2%。     

变弯度后缘与常规舵面机翼的颤振主动抑制对比

后缘变弯度机翼的气动弹性建模与稳定性分析日益受到关注。为了探究变弯度后缘相比常规偏转舵面机翼颤振主动抑制的方法与特点，以一个小展弦比后缘变弯度机翼为对象，首先建立结构有限元模型，并引入变弯度后缘变形模态和常规舵面偏转模态，采用亚声速偶极子格网法计算非定常气动力；然后使用基于最小状态法的有理函数拟合进行频域到时域模型的转换，建立两种构型机翼的气动弹性模型，并在建模时考虑了变弯度后缘与常规舵面控制带宽的差异；最后利用线性高斯二次型方法设计控制律进行颤振主动抑制，分析对比两种控制方式的特性差异。结果表明：采用变弯度后缘的闭环系统能够将颤振临界速度提高22%，其提升效果优于常规舵面，所需舵面偏转峰值更小。     

扇翼飞行器翼型附面层控制数值模拟

基于扇翼飞行器翼型特殊的几何形状及流场特性,在原有翼型的弧形槽下方和后缘加装控制阀门,通过调节阀门开启及开启尺寸的大小,利用弧形槽低压涡所产生的吸力对翼型后缘的附面层进行一定的控制,达到增升减阻的效果。通过采用计算流体力学的方法对其机理及阀门开启尺寸的影响进行了详细计算和分析,研究表明当阀门开启的尺寸为10 mm时,修改翼型的最大升力系数、失速迎角及相同迎角下的升力系数和推力系数均大于基本翼型;随着阀门开启尺寸的增大,修改翼型的最大升力系数和失速迎角均减小,但是在失速前,修改翼型在相同迎角下的升力系数大于基本翼型。此方法可以改变先前通过增大横流风扇的转速来提高其气动性能的做法,减小了能量的消耗,增大了整个飞行器的航程,为扇翼飞行器能够早日投入实际运用奠定了一定的理论基础。     

基于后缘小翼的旋翼翼型动态失速控制分析

针对后缘小翼(TEF)的典型运动参数对旋翼翼型动态失速特性的控制进行了研究。发展了一套适用于带有后缘小翼控制的旋翼翼型非定常流动特性模拟的高效、高精度CFD方法。通过求解Poisson方程生成围绕旋翼翼型的黏性贴体和正交网格,为保证后缘小翼附近的网格生成质量,建立了基于翼型点重构的方法来描述后缘小翼的偏转运动;为克服变形网格方法可能导致网格畸变的不足,发展了一套适用于带有后缘小翼控制的旋翼翼型运动嵌套网格方法。基于非定常雷诺平均Navier-Stokes(URANS)方程、双时间法、Spalart-Allmaras(S-A)湍流模型和Roe-Monotone Upwind-centered Scheme for Conservation Laws(Roe-MUSCL)插值格式,发展了旋翼翼型非定常气动特性分析的高精度数值方法,并采用Lower-Upper Symmetric Gauss-Seidel(LU-SGS)隐式时间推进方法及并行技术提高计算效率。以有试验结果验证的HH-02翼型和SC1095翼型为算例,精确捕捉了动态失速状态下的气动力迟滞效应,验证了本文方法的有效性。着重针对SC1095旋翼翼型的动态失速状态开展后缘小翼的控制分析,提出了可以体现翼型升力、阻力及力矩综合特性的关系式Po和Pc,揭示了后缘小翼振荡频率、相位差和偏转幅值对动态失速特性影响的规律。研究结果表明:当后缘小翼偏转的相对运动频率为1.0,且小翼运动规律与翼型振荡规律之间的相位差为0°时,后缘小翼能够更好地抑制翼型动态失速现象;在此状态下,当偏转幅值为10°时,SC1095翼型最大阻力系数和最大力矩系数可以分别降低19%和27%。     

翼型动态失速的数值研究

用不可压缩流动的求解算法，结合WilcoxDC提出的k－ω模式和k－ωSST湍流模式，对翼型的动态失速进行了数值模拟。通过对典型的振荡翼型轻失速和深失速算例的计算结果分析可以看出：（1）绕动态失速翼型的流场结构十分复杂，轻失速和深失速在流动特性上有很大区别。计算结果显示：轻失速主要是由于后缘分离引起，分离涡的影响范围主要是在后缘附近。而深失速则首先形成很大的前缘分离涡，该分离涡在翼型表面上运动，并诱发出二次分离涡，引起翼型升、阻力系数的显著变化。（2）对于动态失速的翼型绕流，k－ωSST湍流模式是较为有效的，计算出的气动力系数迟带曲线变化趋势与实验结果符合得比较好。     

后缘襟翼对直升机旋翼翼型动态失速特性的影响

 针对带后缘襟翼的智能旋翼直升机典型襟翼参数对翼型动态失速特性的影响进行了研究。建立了带后缘襟翼的桨叶动态失速模型,考虑了襟翼与桨叶之间的缝隙和襟翼在运动过程中相对桨叶的凸起,采用计算流体力学(CFD)方法,研究了不同襟翼转轴位置和襟翼与桨叶的缝隙情况下的翼型动态失速特性,探讨了后缘襟翼激励幅值、时长和起始时刻对升力和俯仰力矩系数的影响。研究结果表明:后缘襟翼能够较好地改善翼型动态失速时的气流环境,并减缓动态失速发生;襟翼激励最优幅值在25°附近,最优激励范围在方位角为240°~360°之间;襟翼转轴后移导致襟翼运动时产生的凸起会使襟翼控制效果减弱;襟翼与桨叶的缝隙会影响翼型动态失速特性,但是缝隙的长度(弦长的2%以内)对襟翼控制效果的影响很小。     

可变弯度机翼后缘形态重构光纤监测技术

可变弯度机翼是一种变翼型变体飞行器，在飞行过程中可根据不同的飞行环境自适应调整机翼弯度来提高飞行效率，从而适应复杂多样的任务环境。针对可变弯度机翼后缘形态与偏转角度实时监测需求，研究了一种基于光纤布拉格光栅传感器的机翼后缘形态重构方法。采用数值仿真方法分析可变弯度机翼后缘的形态变化特征，得到可变弯度机翼后缘偏转位移与偏转角度之间关系。给出光纤布拉格光栅传感器布局形式，构建了基于应变和曲率信息递推位移重构原理的机翼后缘形态和偏转角度监测系统。基于光纤布拉格光栅传感器的机翼后缘形态重构相对误差约为6.39%，偏转角度辨识相对误差约为7.47%。研究结果表明，所提方法能够为可变弯度机翼后缘形态感知、姿态自适应调整以及气动外形优化提供技术支撑。     

电控旋翼气动特性建模与风洞试验验证

首先建立了带襟翼翼型的非定常气动力模型,继而基于Peters-He广义动态尾迹理论,考虑襟翼偏转对电控旋翼叶素环境的影响,建立了电控旋翼有限状态尾迹模型;进一步基于Theodorsen理论推导出电控旋翼桨叶挥舞响应与桨叶变距和襟翼操纵量的关系,综合以上建立了电控旋翼气动特性分析模型.以改进型电控旋翼试验系统为平台进行了风洞试验,测量了不同风速、不同襟翼操纵条件下的电控旋翼气动力、桨距、襟翼偏角及旋翼挥舞角的变化情况.理论计算结果与试验数据符合情况良好,验证了所建立的分析模型的正确性,并得出以下结论:旋翼转速一定时,桨叶变距与襟翼操纵基本呈线性关系;旋翼拉力随襟翼总距的增加而逐渐减小,襟翼总距较大时,其实际气动效率略有下降;前飞状态时,襟翼总距操纵会引起桨叶的纵向周期变距.      

Continuous morphing trailing-edge wing concept based on multi-stable nanomaterial

Morphing technology is one of the most effective methods to improve the flight efficiency of aircraft. Traditional control surfaces based morphing method is mature and widely used on current civil and military aircraft, but insufficiently effective for the entire flight envelope. Recent research on morphing wing still faces the challenge that the skin material for morphing should be both deformable and stiff. In this study, a continuous morphing trailing-edge wing with a new multi-stable nano skin material fabricated using surface mechanical attrition treatment technology was proposed and designed. Computational fluid dynamics simulation was used to study the aerodynamic performance of the continuous morphing trailing-edge wing. Results show that the lift coefficient increases with the increase of deflection angle and so does the lift-drag ratio at a small angle of attack. More importantly, compared with the wing using flaps, the continuous morphing trailing-edge wing can reduce drag during the morphing process and its overall aerodynamic performance is improved at a large angle of attack range. Flow field analysis reveals that the continuous morphing method can delay flow separation in some situations.     

Delaying stall of morphing wing by periodic trailing-edge deflection

Morphing wings can improve aircraft performance during different flight phases. Recently research has focused on steady aerodynamic characteristics of the morphing wing with a flexible trailing-edge, and the unsteady aerodynamic and stall characteristics in the deflection process of the morphing wing are worthy further investigation. The effects of the angle of attack and deflection rate on aerodynamic characteristics were examined, and based on the aerodynamic characteristics of the morphing wing, a method was developed to delay stall by using the flexible periodic trailing-edge deflection. The numerical results show that the lift coefficients in the deflection process are smaller than those in the static situation at small angles of attack, and that the higher the deflection rate is, the smaller the lift coefficients will be. On the contrary, at large angles of attack, the lift coefficients are higher than those in the static case, and they become larger with the increase of the deflection rate. Further, the periodic deflection of the flexible trailing-edge with a small deflection amplitude and high deflection rate can increase lift coefficients at the critical stall angle.     

Aerodynamic characteristics of morphing wing with flexible leading-edge

The morphing wing can improve the flight performance during different phases. However, research has been subject to limitations in aerodynamic characteristics of the morphing wing with a flexible leading-edge. The computational fluid dynamic method and dynamic mesh were used to simulate the continuous morphing of the flexible leading-edge. After comparing the steady aerodynamic characteristics of morphing and conventional wings, this study examined the unsteady aerodynamic characteristics of morphing wings with upward and downward deflections of the leading-edge at different frequencies. The numerical results show that for the steady aerodynamic, the leading-edge deflection mainly affects the stall characteristic. The downward deflection of the leading-edge increases the stall angle of attack and nose-down pitching moment. The results are opposite for the upward deflection. For the unsteady aerodynamic, at a small angle of attack, the transient lift coefficient of the upward deflection, growing with the increase of deflection frequency, is larger than that of the static case. The transient lift coefficient of the downward deflection, decreasing with the increase of deflection frequency, is smaller than that of the static case. However, at a large angle of attack, an opposite effect of deflection frequency on the transient lift coefficient was demonstrated. The transient lift coefficient is larger than that of the static case when the leading edge is in the nose-up stage, and lower than that of the static one in the nose-down stage.     

A CFD-based numerical virtual flight simulator and its application in control law design of a maneuverable missile model

A CFD-based Numerical Virtual Flight (NVF) simulator is presented, which integrates an unsteady flow solver on moving hybrid grids, a Rigid-Body Dynamics (RBD) solver and a module of the Flight Control System (FCS). A technique of dynamic hybrid grids is developed to control the active control surfaces with body morphing, with a technique of parallel unstructured dynamic overlapping grids generating proper moving grids over the deflecting control surfaces (e.g. the afterbody rudders of a missile). For the flow/kinematic coupled problems, the 6 Degree-Of-Freedom (DOF) equations are solved by an explicit or implicit method coupled with the URANS CFD solver. The module of the control law is explicitly coupled into the NVF simulator and then improved by the simulation of the pitching maneuver process of a maneuverable missile model. A nonlinear dynamic inversion method is then implemented to design the control law for the pitching process of the maneuverable missile model. Simulations and analysis of the pitching maneuver process are carried out by the NVF simulator to improve the flight control law. Higher control response performance is obtained by adjusting the gain factors and adding an integrator into the control loop.     

后缘小翼对旋翼气动特性的控制机理及参数分析

针对后缘小翼（TEF）的典型运动参数对旋翼气动特性的控制进行了分析研究。为克服变形网格方法可能导致网格畸变的不足,发展了一套适用于前飞状态带后缘小翼旋翼的运动嵌套网格方法。基于非定常雷诺平均Navier-Stokes（URANS）方程、k-ω剪切应力输运（SST）湍流模型和Roe-MUSCL插值格式,采用含LU-SGS隐式推进的双时间方法及并行技术,建立了一套适用于带有后缘小翼控制的旋翼前飞非定常流动特性模拟的高效CFD方法。以带后缘小翼的SMART旋翼为算例,对比了桨叶剖面等效法向力的计算结果,验证了CFD方法的有效性。着重开展了前飞状态旋翼后缘小翼的控制分析,在操纵量不变的情况下,分别研究了后缘小翼偏转幅值、偏转频率、安装位置及宽度等参数对旋翼气动力的影响特性,获得了典型参数对旋翼气动特性的控制规律。进一步研究了配平状态下后缘小翼对旋翼气动特性的参数影响。结果表明：后缘小翼可以充分发挥旋翼在前行侧的升力潜能,同时降低后行侧动态失速过程中旋翼的阻力和扭矩;在相同的旋翼拉力情况下,通过安装后缘小翼可以将旋翼阻力系数和扭矩系数分别降低17%和29%,升阻比提高14%。     

前端襟翼对带涡襟翼的细长翼影响的实验研究

 为改善带涡襟翼的细长翼的升阻特性,在其上附加了前端襟翼和后缘襟翼。通过前端涡(前端襟翼上产生)和前缘涡(涡襟翼上产生)相互作用对气动特性影响的研究,得到能够改善升阻特性的方法。结果表明,涡的相互作用对涡的产生和发展有很大影响,因而影响细长翼气动特性。附加前端襟翼和后缘襟翼是必要的。由此得到既能增升又能减阻的前端襟翼偏转角。     

基于联合射流的翼型动态失速流动控制研究

动态失速控制对于提高翼型气动特性具有重要意义。采用联合射流方法对翼型俯仰动态失速控制进行数值模拟,完成两方面的研究:一是射流关闭时射流通道对动态失速特性的影响,二是射流开启时不同射流动量系数对动态失速控制的影响和分析。结果表明:射流关闭时,射流通道的存在对翼型上仰过程中附着流阶段的气动特性产生不利影响,使得升力系数明显下降,但是对翼型下俯过程中失速分离流阶段的气动特性影响不明显;射流开启后,动态失速特性得到极大改善,迟滞环面积显著减小,升力增加,阻力减小,且阻力和力矩的峰值显著减小,原基准翼型力矩曲线的负阻尼区域消失。