微型轴流风扇叶顶间隙内泄漏流动的数值模拟

运用FLUENT软件对一种微型轴流风扇内部的非粘性三维流场进行数值模拟,获得了风机内部叶顶间隙内流动规律。对叶顶间隙泄漏涡流的形成和发展作出了比较详细的分析,得到随叶尖间隙的增大,泄漏流动形成泄漏涡,旋涡强度随叶顶间隙的增大(从0.5 mm增大到6 mm)迅速增大到最大值,之后叶顶间隙继续增大,泄漏涡范围增大,但强度不再继续增大。经分析还得到叶顶间隙涡产生于吸力面近叶顶区域,并朝相同流道叶片的吸力面向尾缘方向发展。其结论为进一步弄清叶顶泄漏涡流的流动机理以及为微小型轴流风扇的优化设计都提供了依据。

船用燃气轮机压气机自适应机匣处理设计

为了研究一种在高中低3个转速下都可以扩稳的自适应机匣处理方式,以满足舰船对高速追击性能和良好的低速巡航性的双重要求。本文以实验室现有的一台2.5级压气机为研究模型,在一级动叶叶顶机匣处进行自适应机匣处理设计,根据以往的研究经验制定了一个在高度、开口宽度、喷射角度都较为合理的结构参数,在已有一个喷射口和一个抽吸口的自适应机匣处理上,增加一个抽吸口,以研究对压气机性能的影响。研究结果表明:具有2个抽吸口的自适应机匣处理在3个转速下都有扩稳效果,在100%和75%转速时具有相对低转速更好的扩稳效果,达到2.37%和2.38%,峰值效率损失也较低,在1%左右。这种机匣处理的扩稳机理是,通过改变动叶叶顶附近的叶片表面静压分布,从而改变叶顶泄漏涡流的强度和发展传播方向,减小对主流流动的拦截作用;弦长0.4处的抽吸口有效抑制了此处附近区域二次泄漏涡流的产生,达到延缓失速的目的。

大叶尖间隙下叶栅性能数值模拟

利用NUMECA软件对某线性叶栅的三维流场进行了数值模拟,对比研究了小叶尖间隙和大叶尖间隙时叶尖泄漏涡流的形成和发展,探讨了叶尖间隙大小对叶栅流场和气动性能的影响。研究表明,随着叶尖间隙的增大,叶尖泄漏射流发展成为叶尖泄漏涡,涡流范围不断增大,涡流强度增大趋缓,涡流使得气流偏转减小。从小间隙逐渐增大到大间隙,总压损失与叶片载荷均增大,而在大叶尖间隙时,总压损失增加并不显著,叶片载荷增大趋缓。其结论为进一步揭示叶尖间隙涡流的流动机理以及工业燃气轮机的优化设计提供了参考。

NEW TECHNIQUE FOR CONTROLLING TIP LEAKAGE FLOW OF MICRO TURBINE

A new technique for controlling the tip leakage flow of micro turbine is presented.It uses backward vortex generators（BVGs）to control and reduce the tip leakage flow.The gas is driven by the pressure difference between pressure surface and blade tip,so it flows into BVGs from the pressure side of blade and flows out from the blade tip with high speed in the inverse-direction of leakage flow.Numerical simulations of typical turbine cascade in millimeter scale with and without BVGs are performed.The influence of BVGs on the flow fields and the performance of cascade is analyzed.The results show that the mass of leakage flow decreases by 2.8% and the circumferential load increases by 7.7% under typical entrance condition.Moreover,compared with the case without BVGs,the turbine efficiency grows about 1.2 % in case with BVGs.

发动机冷却风扇叶顶间隙区域泄漏流动的数值分析

针对发动机冷却风扇叶顶间隙区域泄漏的复杂流动和间隙气流对气动性能影响的问题,采用数值分析法研究了导流罩和导流环相对运动和静止对叶顶间隙区域流动的影响、设计工况下间隙区域的压力分布及泄漏涡流的流动特征。采用流场可视化方法测出发动机冷却风扇叶顶间隙区域的流速分布,验证了计算流体力学数值仿真结果的正确性。研究结果表明:随着工作静压的降低,当叶顶间隙区域存在的高速泄漏射流流速与导流环壁面气流周向速度的流速差较大时,由于粘性力的作用,气流流向发生改变,泄漏量明显减小。

Numerical Analysis of the Tip Leakage Flow Field in a Transonic Axial Compressor with Circumferential Casing Treatment

This paper reports on numerical investigations aimed at understanding the influence of circumferential casing grooves on the tip leakage flow and its resulting vortical structures.The results and conclusions are based on steady state 3D numerical simulations of the well-known transonic axial compressor NASA Rotor 37 near stall operating conditions.The calculations carried out on the casing treatment configuration reveal an important modification of the vortex topology at the rotor tip clearance.Circumferential grooves limit the expansion of the tip leakage vortex in the direction perpendicular to the blade chord,but generate a set of secondary tip leakage vortices due to the interaction with the leakage mass flow.Finally,a deeper investigation of the tip leakage flow is proposed.

Relationship between the Flow Blockage of Tip Leakage Vortex and its Evolutionary Procedures inside the Rotor Passage of a Subsonic Axial Compressor

The near casing flow fields inside the rotor passage of a 1.5 stage axial compressor with different blade-loading levels and tip gap sizes were measured by using stereoscopic particle image velocimetry(SPIV). Based on a carefully defined blockage extracting method, the variations of blockage parameter inside the blade passage were analyzed. It was found that the variation of blockage parameter appeared as a non-monotonic behavior inside the blade passage in most cases. This non-monotonic behavior became much more remarkable as the blade loading increases or mass flow rate decreases.The variations of the blockage parameter inside the blade passage had close relation to the evolutionary procedures of the tip leakage vortex(TLV). The destabilization of the TLV caused a rapid increasing of the blockage parameter. After the TLV lost the features of a concentrated streamwise vortex,the blockage parameter usually got a peak value. And then, because of the intense turbulent mixing between the TLV low momentum flow and its surrounding flows, the flow deficit inside the TLV recovered.

Aero-Thermal Performances of Leakage Flows Injection from the Endwall Slot in Linear Cascade of High-Pressure Turbine

The existence of a gap between combustor and turbine endwall in the real gas turbine induces to the leakages phenomenon. However, the leakages could be used as a coolant to protect the endwaU surfaces from the hot gas since it could not be completely prevented. Thus, present study investigated the potential of leakage flows as a function of film cooling. In present study, the flow field at the downstream of high-pressure turbine blade has been investigated by 5-holes pitot tube. This is to reveal the aerodynamic performances under the influenced of leakage flows while the temperature measurement was conducted by thermoehromic liquid crystal （TLC）. Expe- rimental has significantly captured theaerodynamics effect of leakage flows near the blade downstream. Further- more, TLC measurement illustrated that the film cooling effectiveness contours were strongly influenced by the secondary flows behavior on the endwall region. Aero-thermal results were validated by the numerical simulation adopted by commercial sottware, ANSYS CFX 13. Both experimental and numerical simulation indicated almost similar trendinaero and also thermal behavior as the amount of leakage flows increases.

Numerical Investigation of the Interaction between Mainstream and Tip Shroud Leakage Flow in a 2-Stage Low Pressure Turbine

The pressing demand for future advanced gas turbine requires to identify the losses in a turbine and to understand the physical mechanisms producing them. In low pressure turbines with shrouded blades, a large portion of these losses is generated by tip shroud leakage flow and associated interaction. For this reason, shroud leakage losses are generally grouped into the losses of leakage flow itself and the losses caused by the interaction between leakage flow and mainstream. In order to evaluate the influence of shroud leakage flow and related losses on turbine performance, computational investigations for a 2-stage low pressure turbine is presented and discussed in this paper. Three dimensional steady multistage calculations using mixing plane approach were performed including detailed tip shroud geometry. Results showed that turbines with shrouded blades have an obvious advantage over unshrouded ones in terms of aerodynamic performance. A loss mechanism breakdown analysis demonstrated that the leakage loss is the main contributor in the first stage while mixing loss dominates in the second stage. Due to the blade-to-blade pressure gradient, both inlet and exit cavity present non-uniform leakage injection and extraction. The flow in the exit cavity is filled with cavity vortex, leakage jet attached to the cavity wall and recirculation zone induced by main flow ingestion. Furthermore, radial gap and exit cavity size of tip shroud have a major effect on the yaw angle near the tip region in the main flow. Therefore, a full calculation of shroud leakage flow is necessary in turbine performance analysis and the shroud geometric features need to be considered during turbine design process.