超临界翼型Gurney襟翼增升技术实验研究

通过在二元翼型风洞中进行测力实验,研究了不同高度Gurney襟翼对超临界翼型气动力和力矩的影响规律。实验结果表明：在亚声速条件下,Gurney襟翼同样可以明显增加翼型的升力系数,使整个升力曲线向上平移,并使翼型低头力矩增加。高度为翼型弦长0.5%的Gurney襟翼可以带来超临界翼型的最大升阻比。同Gur-ney襟翼对NACA 0012翼型气动特性改变的对比表明,其在超临界翼型上带来的升力系数增量要大于在NACA0012翼型上的效果,但是带来的低头力矩增量较小。     

平板/锯齿型Gurney襟翼对NACA0012翼型增升实验研究

 在Re 为211@10⁶ 情况下进行的NACA0012 翼型Gurney 襟翼增升效应风洞实验研究表明, Gurney襟翼可使升力有很大提高, 0.5%平均气动弦长襟翼在C_L> 11 0 后即可提供较高的升阻比, 当C_L = 11 35 时,2%平均气动弦长襟翼获得了35%的最大升阻比增量; 翼型表面压力分布结果显示, Gur ney 襟翼增加了上翼面的吸力, 同时下翼面压力增强, 因而升力提高; 尾流速度型显示Gur ney 襟翼导致流经上翼面的流体在其后有明显下偏转, 这表明翼型有效弯度增大了; 襟翼上开出锯齿会同时导致升力和阻力下降, 但升阻比是否会提高则应视其是否更接近最佳高度的有效迎风面积。Gur ney 襟翼的最佳应用场合为中高升力系数情况( 如起飞、降落等), 在中小升力系数情况下不宜使用。     

基于涡流发生器和Gurney襟翼的翼型绕流流动控制试验研究

为了研究翼型扰流流动控制,设计了一种三角形涡流发生器(VG)和两种Gurney襟翼,通过Gurney襟翼与VG 组合布局的方式进行风洞试验,对比了干净翼、干净翼加Gurney襟翼、干净翼加VG及干净翼加Gurney襟翼和VG这4种状态下的试验结果.试验结果表明:安装Gurney襟翼对翼型线性段升力在同一迎角下有明显提高...     

Gurney襟翼对翼型流场影响的数值模拟及研究

本文采用计算流体力学的方法计算了NACA23012翼型以及安装四个不同高度（1％C、2％C、3％C、4％C）Gumey襟翼翼型的流场,并比较了不同来流马赫数、气流攻角条件下的气动性能．包括升阻比、翼型表面压力系数分布和马赫数分布。计算结果表明,安装Gumey襟翼后翼型的升阻比得到提高,升力分布趋于平均,在所计算的气流条件下安装3％C高度Gumey襟翼的翼型获得了最高的升阻比。     

Gurney襟翼改善翼型动态失速特性研究

采用CFD数值方法,研究了NACA0012翼型、加装传统Gurney襟翼及改进不对称Gurney襟翼后翼型的动态失速特性。给出了传统襟翼对翼型动态失速特性的影响,并带来较大动态低头力矩的不足,基于传统襟翼的不足,提出了改进的不对称Gurney襟翼方案。研究表明,不对称Gurney襟翼可较好改善翼型的动态失速特性,在增加动态升力的同时,俯仰低头力矩明显减小,可能是直升机旋翼的较理想翼型。     

加装格尼襟翼旋翼的直升机飞行性能

为研究加装格尼襟翼旋翼的直升机飞行性能,建立了加装格尼襟翼旋翼的直升机飞行动力学模型。采用UH-60A直升机试飞数据验证了计算模型的正确性。在此基础上,分析了样例直升机加装格尼襟翼后重量系数、格尼襟翼高度、沿径向位置和加装方式对旋翼需用功率的影响,以及加装格尼襟翼后旋翼桨叶剖面迎角分布、旋翼操纵量和机身姿态角的变化等。研究表明,直升机在重量系数较大的状态下高速前飞时,旋翼加装格尼襟翼能够明显降低直升机的需用功率,且加装转动格尼襟翼的效果优于加装固定格尼襟翼。功率降低幅值随格尼襟翼高度的增加先增加后减小。格尼襟翼在桨叶上布置的位置越靠近桨尖,其对需用功率的影响越大。直升机在重量系数较大的状态下高速前飞时,加装格尼襟翼能够使旋翼后行侧最大迎角显著减小。加装格尼襟翼后旋翼总距和纵横向周期变距减小。     

Gurney襟翼对翼型气动特性影响的风洞实验研究

针对风力机叶片翼型WA251A在Re=3.0×106情况下,安装Gurney襟翼时气动力的变化展开风洞实验研究.对WA251A翼型在干净翼和安装Gurney襟翼情形下的气动力进行对比,重点讨论了Gurney襟翼安装高度、安装角度及安装位置等因素对翼型气动力的影响.实验结果表明:同一安装位置安装高度的增加其增升效应越明显...     

具有Gurney襟翼的多段翼型空气动力特性分析

增大飞机的升力可以有效地缩短飞机起飞和着陆的滑跑距离 ,本文通过对高升力多段翼型有、无Gurney襟翼时的翼面边界层、尾迹速度分布及表面压力分布的测量等实验方法研究了具有Gurney襟翼时的多段翼型绕流特性及增升规律。实验研究结果表明 ,在α =8°时 ,Gurney襟翼高度为 0 .0 2c和 0 .0 5 5c时 ,使多段翼型升力系数分别增加了 1 3%和 2 2 %。Gurney襟翼的增升效果不仅与Gurney襟翼的高度密切相关 ,而且还与在翼面上的安装位置有关。     

Gurney襟翼对某型客机流动控制数值模拟

为改善某型客机的起降性能,通过在机翼尾缘加装Gurney襟翼,对流场进行了数值模拟。对该客机机翼的控制翼型安装不同高度的Gurney襟翼进行数值模拟,结果表明安装Gurney襟翼可以提高多段翼型的升力系数和阻力系数,但会增强尾迹流动的不稳定性。将不同高度的Gurney襟翼应用于该客机的简化模型,机翼的大部分区域符合二维翼型研究得出的流动控制规律;在机翼外侧区域,Gurney襟翼使机翼附近流场中的翼尖涡发生了一定的变化。数值模拟的结果还表明,Gunney襟翼可以提高客机的升力系数,而且不会给飞机流场带来明显的改变。     

Gurney襟翼对双三角翼气动特性影响的低速风洞实验研究

通过低速风洞实验研究了Gurney襟翼对双三角翼气动特性的影响,结果表明Gurney襟翼可以提高双三角翼的升力系数、最大升力系数以及中高升力系数情况下的升阻比。此外,进一步证实了Gurney襟翼的有效迎风面积是影响增升效果的主要因素。     

加装格尼襟翼的自转旋翼气动特性研究

为了研究格尼襟翼对自转旋翼气动特性的影响,首先建立了翼型加装格尼襟翼的二维气动特性计算模型,分析了NACA0012翼型及该翼型加装1%、2%弦长高度格尼襟翼的气动特性,理论计算结果与试验结果的对比表明了本计算模型的正确性。基于叶素理论建立了自转旋翼动力学模型,采用Pitt-Peters动态入流模型捕捉自转旋翼诱导速度沿桨盘的非均匀分布特性。最后进行了自转旋翼加装不同高度格尼襟翼的气动特性分析,结果表明:翼型加装1%弦长高度的格尼襟翼后,在20 m/s到35 m/s的来流速度下,自转旋翼的阻力平均减小可达26%;加装高度为2%弦长的格尼襟翼后,在20 m/s到35 m/s的来流速度下,自转旋翼的阻力平均减小达17%。自转旋翼的气动效率得到明显提高。     

翼型加装格尼襟翼的低雷诺数气动特性实验研究

为了研究低雷诺数下格尼襟翼对翼型气动特性的影响,通过风洞试验研究了Eppler387翼型加装0.5%～5.0%弦长高度格尼襟翼后的气动特性变化,试验雷诺数1.49×105～2.31×105。试验结果表明：低雷诺数下Eppler387翼型加装格尼襟翼后,升力系数和力矩系数明显增大,襟翼高度大于2%弦长时阻力系数显著增大。格尼襟翼在高升力系数下能够起到增大升阻比的作用,适用于微小型飞行器工作在大载荷状态,而0.5%弦长高度的襟翼还能够兼顾中小升力系数下的气动效率,同样适合于微小型飞行器在巡航状态使用。与原翼型相比,加装襟翼后最大升阻比对应的迎角提前,随襟翼高度的增加,翼型升阻比曲线峰值变得不再突出。     

Performance analysis of variable speed tail rotors with Gurney flaps

Gurney Flaps (GFs) are used for improving the performance of variable speed tail rotors. A validated analytical helicopter model able to predict the main and tail rotor power is utilized. The fixed height GF has substantially small influence on the tail rotor power in hover and low to medium speed forward flight, and can obtain significant power reduction in high speed flight. This ability can be enhanced by decreasing the tail rotor speed. With the deployment of GF, the collective pitch of the tail rotor decreases, and the maximum tail rotor thrust increases. The GF can compensate the reduction of the maximum thrust by the decrease in the tail rotor speed. The GF with a height of 5% of the chord length can almost remedy 50% of the thrust reduction introduced by decreasing 10% of the tail rotor speed. With the increase of GF height, the maximum thrust generated by the tail rotor increases. The GF with larger height can cause the increase in the tail rotor power in hover and low to medium speed flight. The retractable GF can obtain more power savings than the fixed height GF. However, the benefit is substantially small even in high speed flight. Considering the side effects introduced by the active GF, the fixed height GF may be more preferable. The mechanism for the retractable GF to generate more tail rotor thrust is to increase the lift in advancing side due to the higher dynamic pressure.     

格尼襟翼在冰风洞混合翼设计中的应用

提出面向工程、面向适航的混合翼设计准则,基于此准则提出将格尼襟翼应用于民机机翼结冰风洞试验的单段混合翼设计中。设计准则中只对驻点位置和前缘吸力峰值提出保守性要求,易于工程实现,满足结冰适航要求。对于偏离设计点较大的状态,在单段混合翼后缘添加格尼襟翼并配合迎角调整,使得单段混合翼在此状态点处仍满足设计准则。这样使得单段混合翼也能应用于试验状态比较广的结冰风洞试验中。格尼襟翼的使用,降低了混合翼设计、试验费用,为民机结冰试航取证节约了时间。     

内吹式襟翼环量控制翼型升力响应特性

传统尖尾缘翼型通过控制迎角,综合利用襟翼、缝翼来改变升力,升力对迎角变化的时间响应历程可以用Wagner函数来描述,而内吹式襟翼（IBF）主要通过控制分离来拓展最大升力,并在一定范围内通过调节射流强度改变驻点位置和环量来对升力进行有效控制,其升力随吹气动量变化的时间响应尺度是否与传统尖尾缘翼型相同还不是很清楚。本文主要研究内吹式襟翼升力响应过程,并将其与传统尖后缘翼型升力响应特性进行对比。首先通过某襟翼偏角为30°的双圆弧环量控制翼型对数值方法进行验证,再对某最大厚度为18%弦长的亚声速翼型内吹式襟翼定常吹气控制下的流场进行非定常数值模拟,并分析了其中的瞬态特征。结果表明内吹式襟翼环量控制翼型对激励响应的时间依赖特征与Wagner函数有很好的相互关系,并可以用该函数来描述。     

Lift enhancement based on virtual aerodynamic shape using a dual synthetic jet actuator

The Dual Synthetic Jet Actuator (DSJA) is used to develop a new type of lift enhancement device based on circulation control, and to control the flow over the two-dimensional (2D) NACA0015 airfoil. The lift enhancement device is composed of a DSJA and a rounded trailing edge (Coanda surface). The two outlets of the DSJA eject two jets (Jet 1 and Jet 2). Jet 1 ejects from the upper trailing edge, which increases the circulation of airfoil with the help of the Coanda surface. Jet 2 ejects from the lower trailing edge, which acts as a virtual flap. The Reynolds number based on the airfoil chord length and free flow velocity is 250000. The results indicate that the circulation control method based on Dual Synthetic Jet (DSJ) has good performance in lift enhancement, whose control effect is closely related to momentum coefficient and reduced frequency. With the increase of the reduced frequency, the control effect of the lift enhancement is slightly reduced. As the momentum coefficient increases, the control effect becomes better. When the angle of attack is greater than 4°, the increments of lift coefficients under the control of DSJ are larger than those under the control of the steady blowing at a same momentum coefficient. The maximum lift augmentation efficiency can reach 47 when the momentum coefficient is 0.02, which is higher than the value in the case with steady blowing jet circulation control.     

Full blended blade and endwall design of a compressor cascade

In the current state-of-the-art, high-loss flow in the endwall significantly influences compressor performance. Therefore, the control of endwall corner separation in compressor blade rows is important to consider. Based on the previous research of the Blended Blade and EndWall (BBEW) technique, which can significantly reduce corner separation, in combination with a non-axisymmetric endwall, the full-BBEW technique is proposed in this study to further reduce the separation in endwall region. The principle of the unchanged axial passage area is considered to derive the geometric method for this technique. Three models are further classified based on different geometric characteristics of this technique: the BBEW model, Inclining-Only EndWall (IOEW) model, and full-BBEW model. The most effective design of each model is then found by performing several optimizations at the design point and related numerical investigations over the entire operational conditions. Compared with the prototype, the total pressure loss coefficient decreases by 7%–9% in the optimized full-BBEW at the design point. Moreover, the aerodynamic blockage coefficient over the entire operational range decreases more than the other models, which shows its positive effect for diffusion. This approach has a larger decrease at negative incidence angles where the intersection of the boundary layer plays an important role in corner separation. The analysis shows that the blended blade profile enlarges the dihedral angle and creates a span-wise pressure gradient to move low momentum fluid towards the mainstream. Furthermore, the inclining hub geometry accelerates the accumulated flow in the corner downstream by increasing the pressure gradient. Overall, though losses in the mainstream grow, especially for large incidences, the full-BBEW technique effectively reduces the separation in corners.     

Modeling and simulating aircraft stability and control—The SimSAC project

This paper overviews the SimSAC Project, Simulating Aircraft Stability And Control Characteristics for Use in Conceptual Design. It reports on the three major tasks: development of design software, validating the software on benchmark tests and applying the software to design exercises. CEASIOM, the Computerized Environment for Aircraft Synthesis and Integrated Optimization Methods, is a framework tool that integrates discipline-specific tools for conceptual design. At this early stage of the design it is very useful to be able to predict the flying and handling qualities of this design. In order to do this, the aerodynamic database needs to be computed for the configuration being studied, which then has to be coupled to the stability and control tools to carry out the analysis. The benchmarks for validation are the F12 windtunnel model of a generic long-range airliner and the TCR windtunnel model of a sonic-cruise passenger transport concept. The design, simulate and evaluate (DSE) exercise demonstrates how the software works as a design tool. The exercise begins with a design specification and uses conventional design methods to prescribe a baseline configuration. Then CEASIOM improves upon this baseline by analyzing its flying and handling qualities. Six such exercises are presented.     

Aerodynamic performance enhancement of a flying wing using nanosecond pulsed DBD plasma actuator

Experimental investigation of aerodynamic control on a 35 swept flying wing by means of nanosecond dielectric barrier discharge(NS-DBD) plasma was carried out at subsonic flow speed of 20–40 m/s, corresponding to Reynolds number of 3.1 · 105–6.2 · 105. In control condition, the plasma actuator was installed symmetrically on the leading edge of the wing. Lift coefficient, drag coefficient, lift-to-drag ratio and pitching moment coefficient were tested with and without control for a range of angles of attack. The tested results indicate that an increase of 14.5% in maximum lift coefficient, a decrease of 34.2% in drag coefficient, an increase of 22.4% in maximum lift-to-drag ratio and an increase of 2 at stall angle of attack could be achieved compared with the baseline case. The effects of pulsed frequency, amplitude and chord Reynolds number were also investigated.And the results revealed that control efficiency demonstrated strong dependence on pulsed frequency. Moreover, the results of pitching moment coefficient indicated that the breakdown of leading edge vortices could be delayed by plasma actuator at low pulsed frequencies.