Influence of Reynolds Number on the Unsteady Aerodynamics of Integrated Aggressive Intermediate Turbine Duct

The ultra-high bypass ratio turbofan engine attracts more and more attention in modern commercial engine due to advantages of high efficiency and low Specific Fuel Consumption (SFC). One of the characteristics of ultra-high bypass ratio turbofan is the intermediate turbine duct which guides the flow leaving high pressure turbine (HPT) to low pressure turbine (LPT) at a larger diameter, and this kind of design will lead to aggressive intermediate turbine duct (AITD) design concept. Thus, it is important to design the AITD without any severe loss. From the unsteady flow’s point of view, in actual operating conditions, the incoming wake generated by HPT is unsteady which will take influence on boundary layer’s transition within the ITD and LPT. In this paper, the three-dimensional unsteady aerodynamics of an AITD taken from a real engine is studied. The results of fully unsteady three-dimensional numerical simulations, performed with ANSYS-CFX (RANS simulation with transitional model), are critically evaluated against experimental data. After validation of the numerical model, the physical mechanisms inside the flow channel are analyzed, with an aim to quantify the sensitivities of different Reynolds number effect on both the ITD and LPT nozzle. Some general physical mechanisms can be recognized in the unsteady environment. It is recognized that wake characteristics plays a crucial role on the loss within both the ITD and LPT nozzle section, determining both time-averaged and time-resolved characteristics of the flow field. Meanwhile, particular attention needs to be paid to the unsteady effect on the boundary layer of LPT nozzle’s suction side surface.     

Numerical investigation of the effects of ITD length on low pressure nozzle

The advantage of high efficiency,low SFC (Specific Fuel Consumption),ultra-high bypass ratio turbofan engine attracts more and more attention in modem commercial engine.The intermediate turbine duct (ITD),which connects high pressure turbine (HPT) with low pressure turbine (LPT),has a critical impact on the overall performance of turbine by guiding flow coming from HPT to LPT without too much loss.Therefore,it becomes more and more urgent to master the technique of designing aggressive,even super-aggressive ITD.Much more concerns have been concentrated on the duct.However,in order to further improve turbine,LPT nozzle is arranged into ITD to shorten low pressure axle.With such design concept,it is obvious that LPT nozzle flow field is easily influenced by upstream duct structure,but receives much less interests on the contrary.In this paper,numerical method is used to investigate the effects of length of ITD with upstream swirl blades on LPT nozzle.Nine models with the same swirl and nozzle blades,while the length of ITD is the only parameter to be changed,will be discussed.Finally,several conclusions and advices for designers are summarized.After changing axial length of ducts,inlet and outlet flow field of nozzle differs,correspondingly.On the other hand,the shearing stress on nozzle blade (suction and pressure) surface presents individual feature under various inlet flow.In addition to that,"Clocking-like effect" is described in this paper,which will contribute much to the pressure loss and should be paid enough attention.     

Numerical investigation of the effects of rising angle on intermediate turbine duct and nearby turbines

In order to improve the efficiency,ultra-high bypass ratio engine attracts more and more attention because of its huge advantage,which has larger diameter low pressure turbine (LPT).This trend will lead to aggressive (high diffusion) intermediate turbine duct (ITD) design.It is necessary to guide the flow leaving high pressure turbine (HPT) to LPT at a larger diameter without any severe loss generating separation or flow disturbances.In this paper,eight ITDs with upstream swirl vanes and downstream LPT nozzle are investigated with the aid of numerical method.These models are modified from a unique ITD prototype,which comes from a real engine.Key parameters like area ratio,inlet height,and non-dimensional length of the ITDs are kept unchanged,while the rising angle (radial offset) is the only changed parameter which ranges from 8 degrees to 45 degrees.In this paper,the effects of rising angle (RA) on ITD,as well as nearby turbines,will be analyzed in detail.According to the investigation results,RA could be as large as 40 degrees in such model of this paper to escape separation;When RA increases,local inlet flow field of LPT nozzle appears to be with apparent variation;while a positive result is that outlet flow field could be kept almost unchanged through modifying blade profile.On the other hand,it seems optimistic that the overall total pressure loss could be kept nearly equivalent among different RA cases.And a valuable conclusion is that outer wall curvature is more important for pressure loss,which advises a clear direction for optimizing ITD.     

Interaction between struts and swirl flow in gas turbine exhaust diffusers

The increasing use of gas turbines in combined cycle power plants together with the high amount of kinetic energy in modern gas turbine exhaust flows focuses attention on the design of gas turbine diffusers as the connecting part between the Brayton/Joule and the Rankine parts of the combined cycle. A scale model of a typical gas turbine exhaust diffuser is investigated experimentally. The test rig consists of a radial type, variable swirl generator which provides the exhaust flow corresponding to different gas turbine operating conditions. Static pressure measurements are carried out along the outer diffuser walls and along the hub of the annular part and along the centerline of the conical diffuser. Velocity distributions at several axial positions in the annular and conical diffuser have been measured using a Laser Doppler Velocimeter (LDV). Pressure recovery coefficients and velocity profiles are depicted as a function of diffuser length for several combinations of swirl strength, tip flow and strut geometries. The diffuser without struts achieved a higher pressure recovery than the diffuser with struts at all swirl angle settings. The diffuser with cylindrical struts achieved a higher pressure recovery than the diffuser with profiled struts at all swirl angle settings. Inlet flows with swirl angles over 18?affected the pressure recovery negatively for all strut configurations.     

Heat transfer and aerodynamics of complex shroud leakage flows in a low-pressure turbine

A numerical investigation on over-shroud & inter-shroud leakage flow has been carried out to explore the underneath flow physics at the stage of shrouded Low Pressure(LP) turbine.Compared with the No inter-Shroud gap’s Leakage flow Model(NSLM) and With inter-Shroud gap’s Leakage flow Model(WSLM),the aerodynamic characteristics and the heat transfer performance have been studied.Through the aerodynamic point of view,it is concluded that due to the pressure difference between the rotor’s passage and the over-shroud cavity,in the stream-wise direction,flow structure has been modified,and the inter-shroud leakage flow may even cause flow separation in the vicinity of the blade passage’s throat.In the circumferential direction,separation flows appear over the rotor’s shroud surface(upper platform of the shroud).Meanwhile,from the point of view of heat transfer,further provision on contour maps of the non-dimensional Nusselt number reveals that the reattachment of leakage flow would enhance the heat transfer rates and endanger the rotor’s labyrinth fins over the shroud.However,due to the limited amount of inter-shroud leakage flow(current computational model),temperature distribution variation along the blade surface(near the rotor’s tip section) seems to have only minor insignificant differences.At the end of the paper,the author puts forward some recommendations for the purpose of future successful turbine design.     

Flowfield and heat transfer past an unshrouded gas turbine blade tip with different shapes

This paper describes the numerical investigations of flow and heat transfer in an unshrouded turbine rotor blade of a heavy duty gas turbine with four tip configurations. By comparing the calculated contours of heat transfer coefficients on the flat tip of the HP turbine rotor blade in the GE-E³ aircraft engine with the corresponding experimental data, the κ-ω turbulence model was chosen for the present numerical simulations. The inlet and outlet boundary conditions for the turbine rotor blade are specified as the real gas turbine, which were obtained from the 3D full stage simulations. The rotor blade and the hub endwall are rotary and the casing is stationary. The influences of tip configurations on the tip leakage flow and blade tip heat transfer were discussed. It’s showed that the different tip configurations changed the leakage flow patterns and the pressure distributions on the suction surface near the blade tip. Compared with the flat tip, the total pressure loss caused by the leakage flow was decreased for the full squealer tip and pressure side squealer tip, while increased for the suction side squealer tip. The suction side squealer tip results in the lowest averaged heat transfer coefficient on the blade tip compared to the other tip configurations.     

叶根倒角对轴流压气机转子两端区流动结构的影响研究

为了揭示叶根倒角对跨音速转子的气动性能影响规律,以NASA的Rotor67转子为研究对象,采用数值方法研究了叶根倒角对跨音速轴流压气机角区分离和工作裕度的影响机制。结果表明：叶根倒角的引入改善了叶片倒角区前缘附近的来流攻角适应性及该区域的叶型几何曲率分布特征,进而提升叶根吸力面的抗分离能力。带有倒角结构的转子叶片在其叶根倒角未覆盖区的叶型中后段周向压力梯度大于原型叶片,有利于克服气流沿吸力面流动时产生的离心力,进而抑制了尾缘附近的分离现象,使得该区域效率提升了3.9%以上。倒角的存在借助于径向平衡约束,重塑了叶尖区域的沿程叶表静压分布,并减小了尖区的入口轴向速度,直接导致叶尖区域主流流体的通流能力明显削弱,并诱发相对更强的间隙泄漏流,从而使得跨音转子提前发生失速,压气机工作裕度降低了19%以上。     

Effects of a kind of non-smooth blade on the unsteady flow field at the exit of an axial fan

An experimental investigation of effects of a kind of streamwise-grooved blade on the unsteady flow field at an exit of an axial-flow fan was performed. The flow field at 25% chord downstream from the trailing edge at hub was measured using a fast-response five-hole pressure probe at different mass-flow conditions. The unsteady flow of the grooved blades was compared with that of the smooth blades. The measurement results indicate that： （1） the grooved blades restrain the velocity fluctuation and the pressure fluctuation by modulating the blade boundary layers, which contributes to the flow loss reduction in the hub region and in the rotor wake region at the design condition; （2） the stream-wise grooves play an important role in restraining the radial migration in the blade boundary layer and abating the tip flow mixing, which contributes to the flow loss reduction in the tip region at the design condition; （3） at the near stall condition, the grooved surface can not reduce the flow loss, even increase the loss nearby when the separation happens in the blade boundary layer.     

舰船燃气轮机变几何动力涡轮通流特性的数值研究

基于叶轮机械三维粘性值模拟技术,本文研究了一个舰船燃气轮机四级变几何动力涡轮的通流特性及其气动性能,并分析了它的主要气流参数随可调导叶转动的变化规律。结果表明,可调导叶级的通流特性将决定整个变几何动力涡轮的气动性能。关小导叶可导致可调导叶级动叶处于较大正攻角而引起吸力面分离流动;反之,开大导叶导致相应的较大负攻角造成压力面分离流动。然而关小导叶引起的吸力面三维分离涡流场将导致变几何动力涡轮的效率更显著的下降。     

Effects of rising angle on upstream blades and intermediate turbine duct

With the improvement of requirement, design and manufacture technology, aero-engines for the future are characterized by further reduction in fuel consumption, cost, but increment in propulsion efficiency, which leads to ultra-high bypass ratio. The intermediate turbine duct (ITD), which connects the high pressure turbine (HPT) with the low pressure turbine (LPT), has a critical impact on the overall performances of such future engines. Therefore, it becomes more and more urgent to master the design technique of aggressive, even super-aggressive ITDs. Over the last years, a lot of research works about the flow mechanism in the diffuser ducts were carried out. Many achievements were reported, but further investigation should be performed. With the aid of numerical method, this paper focuses on the change of performance and flow field of ITD, as well as nearby turbines, brought by rising angle (RA). Eight ITDs with the same area ratio and length, but different RAs ranges from 8 degrees to 45 degrees, are compared.According to the investigation, flow field, especially outlet Ma of swirl blade is influenced by RA under potential effect, which is advisable for designers to modify HPT rotor blades after changing ITD. In addition to that, low velocity area moves towards upstream until the first bend as RA increases, while pressure loss distribution at S2 stream surface shows that hub boundary layer is more sensitive to RA, and casing layer keeps almost constant. On the other hand, the overall total pressure loss could keep nearly equivalent among different RA cases, which implies the importance of optimization.     

1000MW汽轮机低压转子轮槽成形铣刀加工性能研究

采用成形铣刀加工1000MW汽轮机低压转子末级轮槽,通过对26NiCrMoV145转子钢材料切屑的形成情况、刀具磨损形貌以及加工后材料表面残余应力大小的分析,研究成型铣刀的加工性能。结果表明:材料的切屑以带状切屑为主;粗刀以刀刃处边界磨损居多,而半精刀以前刀面损伤更为突出;加工表面残余应力多为残余压应力。     

Investigation of Combined Transpiration and Opposing Jet Cooling of Sintered Metal Porous Struts

Struts are used to inject fuel into the supersonic mainstream of scramjet combustion chambers. The leading edge of the strut experiences the maximum temperature due to tremendous aerodynamic heating. This study describes a combined transpiration and opposing jet cooling system for sintered metal porous struts with a jet flowing out of a micro-slit along the stagnation point line against the incoming flow with methane as the coolant. The combined cooling method for the struts is then compared to cooling by the standard transpiration method. The influences of different slit widths and coolant injection conditions on the strut cooling are numerically investigated. The results show that the combined cooling method significantly reduces the maximum strut temperature. The maximum strut temperature decreases but the coolant consumption increases with increasing micro-slit width. Increasing the micro-slit width more effectively balances the increased cooling effectiveness with the increased coolant flow than just increasing the coolant injection pressure. Coolant injection with non-uniform pressures with higher pressure in the front cavity and lower pressure in the back cavity more effectively enhances the cooling effectiveness and reduces the thermal gradient.     

调节级局部进汽对下游动叶气动负荷影响的数值研究

基于舰船用主汽轮机多工况、变转速、变参数的特点,应用商业软件CFX对具有3组喷嘴的复速级叶栅进行全周数值模拟,着重分析了调节级局部进汽对下游动叶气动负荷的影响,数值模拟结果表明,各个喷嘴组边界壁面的存在对流动具有较大影响,同时喷嘴叶栅Ⅰ的叶片型线的不同导致边界畸变对流动的影响更敏感。动叶叶栅的压力分布特点表明了尾迹低速区和主流低压区是形成动叶流动参数不稳定的主要因素,从而形成了叶栅内特有的非定常流动现象。     

Numerical Investigations of the Influence of Unsteady Vane Trailing Edge Shock Wave on Film Cooling Effectiveness of Rotor Blade Leading Edge

Unsteady numerical simulations of a high-load transonic turbine stage have been carried out to study the influences of vane trailing edge outer-extending shockwave on rotor blade leading edge film cooling performance. The turbine stage used in this paper is composed of a vane section and a rotor one which are both near the root section of a transonic high-load turbine stage. The Mach number is 0.94 at vane outlet, and the relative Mach number is above 1.10 at rotor outlet. Various positions and oblique angles of film cooling holes were investigated in this research. Results show that the cooling efficiency on the blade surface of rotor near leading edge is significantly affected by vane trailing edge outer-extending shockwave in some cases. In the cases that film holes are close to leading edge, cooling performance suffers more from the sweeping vane trailing edge outer-extending shockwave. In addition, coolant flow ejected from oblique film holes is harder to separate from the blade surface of rotor, and can cover more blade area even under the effects of sweeping vane trailing edge shockwave. As a result, oblique film holes can provide better film cooling performance than vertical film holes do near the leading edge on turbine blade which is swept by shockwaves.     

一级半轴流式透平级内部若干非定常流动现象的试验研究

设计研制了具有亚音速透平高压级气动特性的一级半轴流式试验透平,采用试验方法对时序效应、叶栅壁面非定常静压幅频特性以及动叶出口非定常速度场进行了研究。结果表明:时序效应具有改善轴流式透平气动性能的潜力;动、静叶排压力有势场干涉引发的基频信号和上游静叶尾迹片段引发的两阶倍频信号,构成了第二列静叶壁面静压非定常分量的基本频率特征,其间还伴随高达六阶的倍频信号,主要由动叶尾缘高频脱落的涡街扰动产生;尚未完成掺混的第一列静叶尾迹片段出现在动叶出口,由其引发的负射流显著改变了动叶出口局部位置处的气流偏转角。     

Broadband noise radiation analysis for an HAWT rotor

The paper demonstrates a computational methodology for the noise prediction of an horizontal axis wind turbine (HAWT) rotor in time domain. The aeroacoustic modeling is based on the Ffowcs Williams and Hawkings equation considering only the surface monopole and dipole noise sources. The accurate solution of the noise problem is strongly influenced by the unsteadiness of the rotor flowfield, the nonuniform inflow effects and the blade aerodynamic parameters which are included in the numerical model. The aerodynamic analysis uses a three-dimensional low-order panel method combined with a boundary layer correction model for the calculation of the blade pressure distribution, obtaining satisfactory agreement with experimental data. Acoustic pressure predictions for the NREL downwind rotor are presented for several cases proving that broadband noise propagation dominates at low- and mid-frequency ranges. Sound pressure level contours occur via Fourier transformations, while the directivity for a single frequency source is also examined.     

Numerical simulation on the effect of jet nozzle position on impingement cooling of gas turbine blade leading edge

In this paper, Computational Fluid Dynamics (CFD) simulations are performed to investigate the impingement cooling on internal leading edge region which is stretched by the middle cross section of the first stage rotor blade of GE-E³ engine high pressure gas turbine. The simulations are carried out for a blade with a single row of circle jets at five different positions and seven different inlet flow Mach numbers. The results indicate that the global area weighted average Nusselt number at the blade leading edge increases with the increase of jet Mach number, and increases with the decrease of the distance between the jet nozzle and the pressure side. The correlation for the area weighted average Nusselt number as a function of the parameters is derived for the range of the parameters considered. The streamwise length weighted average Nusselt number and the spanwise length weighted average Nusselt number also increase with the decrease of the spacing between the jet nozzle and the pressure side, and increase with the increase of jet Mach number. The side entry jet is desirable to improve the performance of impingement cooling on turbine leading edge, but the arrangement of the jet nozzle and the shape of the internal cooling passage should be further optimized to improve the distribution of the heat transfer coefficient.     

Wave energy harvesting using a novel turbine rotor geometry

Wells turbines provide a practical solution for wave energy harvesting. The low aerodynamic efficiency of Wells turbines tangibly reduces their output power. Both the turbine efficiency and output power depend on the turbine solidity. The turbine solidity decreases from rotor hub to rotor tip for the commonly used rotors with constant chord‐length blades. The present work introduces a novel Wells turbine rotor geometry. This geometry was obtained by numerically optimizing the rotor's radial solidity distribution. The turbine performance with different rotor geometries was numerically simulated by solving the three‐dimensional Reynolds‐averaged Navier–Stocks equation under incompressible and steady state flow conditions. Simple and multi‐objective optimization were implemented in order to obtain the optimum rotor geometry. The present work showed that an improved turbine performance can be achieved by optimizing the turbine radial solidity distribution. Two different optimized rotor geometries were obtained and presented. The first rotor geometry improved the turbine efficiency by up to 4.7% by reducing its pressure drop. The second rotor geometries enhanced the turbine output power by up to 10.8%. Copyright © 2016 John Wiley & Sons, Ltd.     

工况不确定性对汽轮机末两级气动性能的影响研究

汽轮机运行条件存在随机波动性,对汽轮机的稳定运行造成一定影响。以往的研究多将汽轮机的工作环境条件作为确定因素,未考虑其随机变化的影响。本文将汽轮机背压、径向气流角和进口流量作为服从一定概率分布的随机变量,采用多项式混沌方法结合计算流体力学（CFD）模拟仿真,研究了以上参数的随机波动对某汽轮机末两级叶片气动性能的影响。结果表明：计算进口边界条件的设置方法对汽轮机末两级总体性能的计算结果影响不大,当给定进口流量边界条件时,末两级效率的计算结果最高;在接近堵塞工况时,进出口条件的随机变化对通流流量的影响不大,但效率存在明显的波动,进口气流角对总体性能的影响相对较小;当背压和进口流量存在随机波动时,末级动叶中激波位置小范围波动,激波位置对径向气流角的波动不敏感。     

尾气颗粒对可变混流涡轮增压器喷嘴环冲蚀磨损特性的数值模拟

可变涡轮增压器在运行过程中其涡轮喷嘴环会受到尾气颗粒的冲蚀磨损,造成喷嘴环叶片失效和涡轮运行效果下降。借助CFD软件对可变混流涡轮内部的气固两相流进行数值模拟计算,分析喷嘴环开度的变化和颗粒粒径对喷嘴环磨损规律的影响。结果表明：喷嘴环处于不同开度下时,磨损率和磨损区域均有所不同,磨损区域主要集中在喷嘴环压力面的中后段区域,随着开度的增大,磨损率和磨损区域均减小,压力面的磨损程度明显大于吸力面;尾气中的小颗粒因惯性较小对气流的跟随性较好,主要撞击喷嘴环前缘,且开度对喷嘴环前缘的磨损影响较小;尾气中大颗粒的运动轨迹趋于直线,主要撞击喷嘴环压力面的中后段区域;由于涡轮进气涡壳结构的周向不对称性,使得涡轮内部流场沿周向分布不均匀,导致不同周向位置的喷嘴环磨损率和磨损区域有所不同,且随着开度的增大,各喷嘴环之间的磨损差异也增大。