共查询到20条相似文献,搜索用时 31 毫秒
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为改善流动分离造成叶片气动效率降低,基于鸟鹰类翅膀羽毛在大范围流动分离时自适应弹起的特点,在翼型吸力面设置功能类似羽毛的弹片。弹片在未发生大范围流动分离时贴附翼型表面,使原始翼型轮廓发挥作用,并于攻角增大时弹起以改善翼型失速特性。以NREL S809为原始翼型,对不同攻角下多个弹片角度进行了数值计算,并对所得气动参数进行分析。研究表明:在大范围流动分离时,弹片可有效提高升阻比,最高达50%~60%;气流贴附弹片流动至其末端,从而抑制和拖延了涡的发展,进而提高了流场稳定性,使波动更规律且幅度更小;所研究攻角范围内,改善翼型气动性能的最佳弹片角度随攻角呈近似线性变化。 相似文献
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通过研究尾缘气动弹片对翼型动态失速特性影响,提出一种基于气动弹片的主动控制策略,使其于大攻角时抬起,小攻角时闭合。并采用计算流体动力学方法对比分析主动式气动弹片对不同厚度翼型抑制流动分离作用的效果。结果表明:对于薄翼型,发生动态失速时,气动弹片可延缓翼型尾缘涡旋与前缘主流涡的相互作用,减小翼型升力系数骤降幅度;随翼型厚度增加,流动分离点从翼型前缘转向后缘,气动弹片可有效分割较大分离涡,减轻流动分离程度,限制分离涡发展,同时抑制尾缘伴随小涡产生,提高翼型升阻比。 相似文献
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This paper presents a detailed numerical investigation of the influence of re-organized shock waves on the flow separation for a highly-loaded transonic compressor cascade. The boundary layer suction (BLS) was used to control the location and strength of shock waves, with the aspirated slot locating at 49% chord, where is just downstream of the impingement point of shock wave at the leading edge. The numerical simulation is based on NUMECA, a commercial software, where the cell-centered control volume approach with third-order spatial accuracy is used to solve the 3-D Reynolds-averaged Navier-Stokes equations under the Cartesian coordinate system. Several conclusions can be made through the observation of the numerical results. (1) Multiple shock waves in cascade passage leaded the velocity deficits of boundary layer on suction surface downstream of shock wave, resulting in seriously separated flow on the suction side of blade, especially when the front shock wave is much stronger than the rest of the shocks. (2) BLS with small mass flow rate can not effectively improve the boundary layer. When the impingement point of oblique shock wave coming from cascade leading edge is bled to downstream of the passage shock wave by BLS, the boundary layer flow is greatly improved. However, if the BLS mass flow rate exceeds a critical value (1.2%), the boundary layer downstream of shock wave would separate from suction surface. (3) At the blade mid-span, the aerodynamic performance of compressor blade is improved as BLS mass flow rate increases. The optimum BLS is about 1.2%. Compared with the baseline case, the BLS with flow rate of 1.2% increases the total pressure recovery coefficient by 12%, and decreases diffusion factor by 18% and deviation angle to 7 ° while keeping the pressure rise constant. (4) The three dimensional flow structure of the compressor cascade ranged from 25% span to 75% span was improved greatly with the 1.2% BLS flow rate. However it could not control the development of the corner boundary layer effectively. 相似文献
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翼缝是翼型主体与襟翼之间的缝隙,对翼型气动性能与流场结构有很大影响。以两段式NACA0018翼型为基础翼型,对传统弯曲翼缝进行改进设计与数值模拟,以期增大失速攻角及改善在大攻角下的气动性能。结果表明:在小攻角下,导叶翼缝襟翼翼型的升力较原始NACA0018翼型小,阻力较大,但在大攻角下,导叶翼缝可减小翼缝中流体的速度损失,为翼型上表面边界层提供更多动能,从而改善流场结构及失速特性,弯曲翼缝可增大1°失速攻角,而导叶翼缝可增大8°,攻角为18°时升力系数较弯曲翼缝提升43%。因此,导叶翼缝可极大地改善翼型在大攻角下的气动性能。 相似文献
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Efficiency enhancement in transonic compressor rotor blades using synthetic jets: A numerical investigation 总被引:3,自引:0,他引:3
Several passive and active techniques were studied and developed by compressor designers with the aim of improving the aerodynamic behavior of compressor blades by reducing, or even eliminating, flow separation. Fluidic-based methods, in particular, have been investigated for a long time, including both steady and unsteady suction, blowing and oscillating jets. Recently, synthetic jets (zero mass flux) have been proposed as a promising solution to reduce low-momentum fluid regions inside turbomachines. Synthetic jets, with the characteristics of zero net mass flux and non-zero momentum flux, do not require a complex system of pumps and pipes. They could be very efficient because at the suction part of the cycle the low-momentum fluid is sucked into the device, whereas in the blowing part a high-momentum jet accelerates it. To the authors’ knowledge, the use of synthetic jets has never been experimented in transonic compressor rotors, where this technique could be helpful (i) to reduce the thickness and instability of blade suction side boundary layer after the interaction with the shock, and (ii) to delay the arising of the low-momentum region which can take place from the shock-tip clearance vortex interaction at low flow operating conditions, a flow feature which is considered harmful to rotor stability. Therefore, synthetic jets could be helpful to improve both efficiency and stall margin in transonic compressor rotors. In this paper, an accurate and validated CFD model is used to simulate the aerodynamic behavior of a transonic compressor rotor with and without synthetic jets. Four technical solutions were evaluated, different for jet position and velocity, and one was investigated in detail. 相似文献
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Hamzeh Eshraghi Masoud Boroomand Abolghasem M. Tousi Mohammad Toude Fallah Ali Mohammadi 《热科学学报(英文版)》2016,25(3):223-230
Increasing the aerodynamic load on compressor blades helps to obtain a higher pressure ratio in lower rotational speeds.Considering the high aerodynamic load effects and structural concerns in the design process,it is possible to obtain higher pressure ratios compared to conventional compressors.However,it must be noted that imposing higher aerodynamic loads results in higher loss coefficients and deteriorates the overall performance.To avoid the loss increase,the boundary layer quality must be studied carefully over the blade suction surface.Employment of advanced shaped airfoils (like CDAs),slotted blades or other boundary layer control methods has helped the designers to use higher aerodynamic loads on compressor blades.Tandem cascade is a passive boundary layer control method,which is based on using the flow momentum to control the boundary layer on the suction surface and also to avoid the probable separation caused by higher aerodynamic loads.In fact,the front pressure side flow momentum helps to compensate the positive pressure gradient over the aft blade's suction side.Also,in comparison to the single blade stators,tandem variable stators have more degrees of freedom,and this issue increases the possibility of finding enhanced conditions in the compressor off-design performance.In the current study,a 3D design procedure for an axial flow tandem compressor stage has been applied to design a highly loaded stage.Following,this design is numerically investigated using a CFD code and the stage characteristic map is reported.Also,the effect of various stator stagger angles on the compressor performance and especially on the compressor surge margin has been discussed.To validate the CFD method,another known compressor stage is presented and its performance is numerically investigated and the results are compared with available experimental results. 相似文献
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尾缘襟翼对风力机翼型气动特性影响研究 总被引:1,自引:0,他引:1
尾缘襟翼(TEF)因其对翼型气动特性的调控能力,被认为是降低叶片疲劳和局部载荷最具可行性的气动控制部件。对TEF进行建模,采用Xfoil和CFD软件分析了TEF对翼型气动特性的影响及其机理,并从叶素理论角度对变化来流下TEF的减载效果进行了验证,结果表明:TEF位于不同摆角时翼型升阻力系数均有不同程度的变化,TEF可有效实现对翼型气动特性的主动控制;TEF摆动改变了翼型表面的静压分布和流动状态,进而对翼型升阻力和失速攻角产生影响;TEF可快速有效降低风速突然增加后的叶素受力,进而控制并减小叶片载荷。 相似文献
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采用环形亚声速风洞气动参数测量试验方法,研究了某型轴流压气机高压第16级动、静叶型的变攻角特性。试验结果表明,对于扩压叶型,沿压力边和吸力边的流动都是先快速膨胀后扩压,扩压占流程的绝大部分,相对压力边,吸力边扩压流程长,梯度大,吸力边是叶型损失源;在相同攻角下,叶型损失正比于叶型的几何折转角,因为在试验攻角变化范围内,叶型的落后角可忽略不计。 相似文献
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以Phase Ⅵ风力机叶片为研究对象,以r/R=30%、63%和95%处叶素为参考,建立与7、9、15 m/s试验风速下该风力机叶片附着涡环量沿展向分布相同的叶片模型,分析尾随涡对风力机当地翼型气动性能的影响机理。采用带转捩效应的SST k-ω湍流模型,对所建立的叶片模型和二维S809翼型的气动特性进行研究和对比分析。结果表明:旋转叶片尾随涡对分离现象产生抑制作用且随攻角的增大减弱;尾随涡的影响表现出多重效应,除了减小当地翼型的攻角,还降低其吸力面负压系数和压力面正压系数。 相似文献
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基于定常RANS方程,采用Spalart-Allmaras(S-A)湍流模型,数值模拟某跨音速导叶尾缘劈缝射流的定常流动结构,分析尾缘劈缝射流对尾缘激波结构、尾迹流动特性及叶栅气动性能的影响。研究表明:开缝射流显著降低尾缘压力面侧燕尾波强度,并使激波在相邻叶片吸力面入射点向上游移动;当叶栅出口马赫数小于1.35时射流使吸力面燕尾波强度减弱,而达到1.35后射流使该侧激波强度增大;在不同出口马赫数下射流均能降低叶栅动能损失。 相似文献
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Soichi Sasaki Hajime Takamatsu Masao Tsujino Haruhiro Tsubota Hidechito Hayashi 《热科学学报(英文版)》2010,19(1):60-66
In order to clarify the mechanism by which aerodynamic noise is generated from separated flow around an airfoil blade, the relation between the attack angle and the aerodynamic noise of the blade was analyzed using a wind tunnel experiment and a CFD code. In the case of rear surface separation, the separated vortex which has a large-scale structure in the direction of the blade chord is transformed into a structure that concentrates at the trailing edge with an increase in the attack angle. The aerodynamic noise level then becomes small according to the vortex scale in the blade chord. When the flow is separated at the leading edge, a separated vortex of low pressure is formed at the vicinity of the trailing edge. The pressure fluctuations on the blade surface at the vicinity of the trailing edge become large due to the vortex in the wake. It is considered that the aerodynamic noise level increases when the flow is separated at the leading edge because the separated vortex is causing the fluctuations due to wake vortex shedding. 相似文献
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为研究间隙变化对轴流压气机转子近失速工况下叶顶流场结构的影响,以轴流压气机转子Rotor37为研究对象,对其叶顶流场进行定常和非定常的数值模拟。计算结果表明:随着叶顶间隙的减小,压气机的总压比和等熵效率均有所提高,稳定运行范围扩大;2倍设计间隙下,叶尖泄漏涡经激波作用后发生膨胀破碎,堵塞来流通道,诱发压气机堵塞失速;0.5倍设计间隙下,吸力面流动分离加剧,发生回流,部分回流与来流在压力面前缘上游发生干涉,进口堵塞加剧,致使部分来流从前缘溢出,导致压气机叶尖失速;不同间隙下压气机失速过程的主导因素不同,大间隙下失速由叶尖泄漏涡破碎的非定常波动引起,小间隙下失速主要由流动分离引发的周期性前缘溢流所主导。 相似文献
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以NACA0018翼型为原始模型进行前缘结构设计,采用计算流体动力学(CFD)方法分析凹凸前缘结构参数对叶片绕流流动及气动性能的影响。结果表明:在0°~10°攻角范围内,凹凸前缘叶片气动性能与原始叶片基本一致,但在15°~25°攻角范围内,正弦波形凹凸前缘叶片升力系数最大提升20.2%;叠加波形凹凸前缘叶片在15°~25°攻角内,气动性能均有不同程度的下降,波峰处推迟分离,而在波谷分离提前,在吸力面每个波谷顺流方向叶片及展向形成反向涡对,相互卷吸并与主流掺混增加能量交换向尾缘处移动,改变了叶片原始流场反馈回路,阻碍了叶片展向涡及流向涡的发展。 相似文献