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.     

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.     

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.     

Effects of nozzle-strut integrated design concepton on the subsonic turbine stage flowfield

In order to shorten aero-engine axial length,substituting the traditional long chord thick strut design accompanied with the traditional low pressure（LP） stage nozzle,LP turbine is integrated with intermediate turbine duct（ITD）.In the current paper,five vanes of the first stage LP turbine nozzle is replaced with loaded struts for supporting the engine shaft,and providing oil pipes circumferentially which fulfilled the areo-engine structure requirement.However,their bulky geometric size represents a more effective obstacle to flow from high pressure（HP） turbine rotor.These five struts give obvious influence for not only the LP turbine nozzle but also the flowfield within the ITD,and hence cause higher loss.Numerical investigation has been undertaken to observe the influence of the Nozzle-Strut integrated design concept on the flowfield within the ITD and the nearby nozzle blades.According to the computational results,three main conclusions are finally obtained.Firstly,a noticeable low speed area is formed near the strut＇s leading edge,which is no doubt caused by the potential flow effects.Secondly,more severe radial migration of boundary layer flow adjacent to the strut＇s pressure side have been found near the nozzle＇s trailing edge.Such boundary layer migration is obvious,especially close to the shroud domain.Meanwhile,radial pressure gradient aggravates this phenomenon.Thirdly,velocity distribution along the strut＇s pressure side on nozzle＇s suction surface differs,which means loading variation of the nozzle.And it will no doubt cause nonuniform flowfield faced by the downstream rotor blade.     

某MW级燃机非对称排气道性能分析

针对某MW级燃机动力透平及非对称排气道,本文采用CFD方法首先单独研究了不同进口条件下排气道的性能,发现该排气道性能良好,总压损失较低,且对15°以内的入口旋流角具有较好的适应性。由于动力透平与排气道之间存在相互作用,将整周动力透平与排气道进行耦合计算,发现由于动力透平设计存在一定的缺陷,单级计算时动叶根部存在大范围的气流分离,与排气道耦合后,透平功率略有下降,但动叶根部气流展向分离范围有所减小,且分离程度出现周向不均匀现象;同时,由于气流分离导致动力透平出口气流的旋流角较大,耦合条件下排气道性能明显恶化,总压损失增加较大。     

发动机诱导产生进气涡流和可调进气涡流的研究

在稳流气道试验台上开展了诱导产生发动机进气涡流和可调进气涡流的研究工作,证明在短直气道内,通过安装进气导流片能够产生一定强度的进气涡流,再配合另一气道进气流量的控制,可获得发动机缸内可调的进气涡流。非增压发动机试验表明在低速低负荷工况下,发动机的燃油消耗率有明显的改善。     

Study over thermal state of gas turbine engine metal-ceramic rotor blades and nozzle guide vanes under thermal shock and thermal-cyclic loading conditions

To ensure a reliable operation of the 2.5 MW gas turbine engine (GTE- 2.5) with the inlet gas temperature TIT = 1623 K, studies were performed over the thermal state of the nozzle guide vanes and rotor blades with the temperatures, rates and flows of the working media and cooling air simulating all the potential turbine stage operating duties. The steady state and thermal-cyclic tests having been accomplished, there was no visible defect on the rotor blades and the nozzle vanes. Afterwards, they survived the endurance tests at the rated cooling. Therefore, the functionality of the shell thin-wall hybrid nozzle vanes and rotor blades under the variable operating duties of the gas turbine at the "shock" and "cyclic" loads of the working media temperature variations has been demonstrated.     

A method for the determination of turbulence intensity by means of a fast response pressure probe and its application in a LP turbine

This paper describes the measurements and the post-processing procedure adopted for the determination of the turbulence intensity in a low pressure turbine (LPT) by means of a single sensor fast response aerodynamic pressure probe. The rig was designed in cooperation with MTU Aero Engines and considerable efforts were put into the adjustment of all relevant model parameters. Blade count ratio, airfoil aspect ratio, reduced massflow, reduced speed, inlet turbulence intensity and Reynolds numbers were chosen to reproduce the full scale LP turbine. Measurements were performed adopting a phase-locked acquisition technique in order to provide the time resolved flow field downstream of the turbine rotor. The total pressure random fluctuations are obtained by selectively filtering, in the frequency domain, the deterministic unsteadiness due to the rotor blades and coherent structures. The turbulence intensity is derived from the inverse Fourier transform and the correlations between total pressure and velocity fluctuations. The determination of the turbulence intensity allows the discussion of the interaction processes between the stator and rotor for engine-representative operating conditions of the turbine.     

可控进气涡流对柴油机低负荷性能影响的研究

就双进气道可控进气涡流系统改善大功率增压中冷柴油机低负荷性能的作用机理进行了理论分析与试验研究。研究结果表明：在低负荷工况下，双气道进气时的涡流强度小于螺旋气道，大于直流气道；螺旋气道进气流量较之双气道虽有所减少，但其涡流比仍保持较高水平，致使缸内爆压、预混合燃烧量和最大放热速率均较大。理论计算与试验结果证明，采用双进气道可控进气涡流系统是改善大功率柴油机低负荷性能的一个有效途径。     

发动机氧涡轮喷嘴叶栅气动特性的试验研究

对一个用于大推力液体火箭发动机氧涡轮泵的复速级涡轮的喷嘴叶栅进行了试验研究，以考察喷嘴叶栅的气动特性，验证喷嘴叶栅的气体设计。该复速级喷嘴叶栅采用先进的后加载流动控制技术，以减弱叶机的二次流损失，对喷嘴叶栅进行了四个进气口流角，三个出口等熵马赫数条件下的平面叶栅吹风试验，测取了型面压力分布，出口气流角以及叶栅损失等重要气动特性参数，试验研究表明氧涡轮的喷嘴叶栅的设计是成功的，具有良好的气动特性，可以有效地应用于液体火箭发动机的涡轮中，本研究也为该类喷雾叶栅的设计提供了有用的实验数据和指导意义的结论。     

变几何低压涡轮级多工况气动性能研究

为明确变几何低压涡轮级在多转角工况下气动性能变化情况,通过RANS方法并结合SST湍流模型,研究了可调导叶转角分别为-6°,-3°,0°,3°和6°条件下低压涡轮级的气动性能变化。结果表明：可调导叶旋转角度的变化会明显改变导叶叶顶及动叶通道内的流动情况,角度变大会增加涡轮级流量,并使导叶叶顶处负荷后移,上端区二次流强度增加,叶顶泄漏情况减弱,还会减小动叶进口相对气流角,使动叶压力面出现明显分离;角度变小对低压涡轮级流场的影响与之相反。当导叶转角从-6°变化到+3°时,涡轮级等熵滞止效率提升了约6.7%;当导叶转角从+3°变化到+6°时,涡轮级效率却下降了约0.19%。     

入口旋流角对排气扩压段气动性能影响的数值研究

为了研究入口气流旋流角对带支撑结构轴流排气扩压段气动性能的影响,以某型燃气轮机排气扩压段为研究对象,采用数值计算的方法,对单独排气段模型及排气段和涡轮末级动叶耦合模型分别进行数值模拟。采用总压保持系数和静压恢复系数作为衡量排气扩压段气动性能的主要参数。排气段单独模拟的结果显示,当旋流角从0°变化至-32°,总压保持系数下降4%,且在-20°以后开始呈现突然的快速下降趋势;而静压恢复系数先上升后降低,在-16°时达到最大值。另外,通过耦合模型与单独排气段模型的数值计算对比,发现当排气段入口旋流角和质量流量相同时,计算结果较为一致。以上结果说明,入口旋流角是影响排气扩压段流动性能的关键因素之一,进行排气段结构设计时要充分考虑旋流角对内部流动的影响。而且,单独排气段数值模拟在相同质量流量和旋流角度条件下,可近似达到耦合模拟的精度,提高设计效率。     

6FA燃气轮机进气系统流量校准方案分析

本文基于某6FA燃气轮机项目进气喷嘴校准装置的设计过程,通过建立进气道模型,选择与安装流量喷嘴、测压耙与扫描阀的方案,说明安装进气流量校准系统的作用与安装校准装置对进气流量喷嘴空气流场的影响,结果可帮助提高燃气轮机进气流量测量精度。     

Numerical study of combining steady Vortex Generator Jets and Deflected Trailing edge to reduce the blade numbers of low pressure turbine stage

A flow control system that combined steady Vortex Generator Jets and Deflected Trailing-edge (VGJs-DT) to decrease the low pressure turbine (LPT) blade numbers was presented. The effects of VGJs-DT on energy loss and flow of low solidity low pressure turbine (LSLPT) cascades were studied. VGJs-DT was found to decrease the energy loss of LSLPT cascade and increase the flow turning angle. VGJs-DT decreased the solidity by 12.5% without a significant increase in energy loss. VGJs-DT was more effective than steady VGJs. VGJs-DT decreased the energy loss and increased the flow angle of the LSLPT cascade with steady VGJs. VGJs-DT can use 50% less mass flow than steady VGJs to inhibit the flow separation in the LSLPT cascade. The deflected trailing edge enhanced the ability of steady VGJs to resist flow separation. Overall, VGJs-DT can be used to control flow separation in LPT cascade and reduce the blade numbers of low pressure turbine stage.     

Fundamental Mechanisms That Influence the Estimate of Heat Transfer to Gas Turbine Blades

The quest for improved efficiency has motivated the elevation of turbine inlet temperatures in all types of advanced aircraft gas turbines. The accommodation of higher gas temperatures necessitates complex blade cooling schemes so as not to sacrifice structural integrity and operational life in advanced engine designs. Estimates of the heat transfer from the gas to stationary (vanes) or rotating blades poses a major uncertainty because of the complexity of the heat transfer processes. The gas flow through these blade rows is three-dimensional with complex secondary viscous flow patterns that interact with the end walls and blade surfaces. In addition, upstream disturbances, stagnation flow, curvature effects, and flow acceleration complicate the thermal transport mechanisms in the boundary layers. Some of these fundamental heat transfer effects will be discussed. The chief purpose of this paper is to acquaint those in the heat transfer community, who are not directly involved in gas turbines, with the seriousness of the problem and to recommend some basic research that would improve the predictions of gas-side heat transfer on turbine blades and vanes.     

Influences of inflow condition on non-axisymmetric flows in turbine exhaust hoods

The complex 3D flow in a steam turbine exhaust hood model with different inlet swirl and inlet total pressure radial distributions has been simulated by employing CFX-5 and analyzed in this paper. It＇s found that the inlet tangential flow angle at hub has a negative effect on the exhaust hood performance, while a negative gradient of inlet total pressure radial distribution has a positive impact on the hood performances. It＇s also numerically con- firmed that a proper distribution of total pressure at hood inlet can successfully eliminate the negative effects caused by the inappropriate inlet swirl distribution and improve the hood aerodynamic performance.     

Effects of Variable Jet Nozzle Angles on Cross-Flow Suppression and Heat Transfer Enhancement of Swirl Chamber

This paper investigated the effects of variable jetting nozzle angles on the cross-flow suppression and heat transfer enhancement of swirl cooling in gas turbine leading edge. The swirl chamber with vertical jet nozzles was set as the baseline, and its flow fields and heat transfer characteristics were analyzed by 3D steady state Reynolds-averaged numerical methods to reveal the mechanism of cross-flow weakening the downstream jets and heat transfer. On this basis, the flow structure on different cross sections and heat transfer characteristics of swirl chamber with variable jetting nozzle angels were compared with the baseline swirl chamber. The results indicated that for the baseline swirl chamber the circumferential velocity gradually decreased and the axial velocity gradually increased, and the cross-flow gradually formed. The cross-flow deflected the downstream jets and drawn them to the center of the chamber, thus weakening the heat transfer. For swirl chamber with variable jetting nozzle angles, the air axial velocity is axial upstream, opposite to the mainstream, so that the impact effects of cross-flow on the jets were reduced, and the heat transfer was enhanced. Furthermore, with the increase of axial velocity along the swirl chamber, the jetting nozzle angle also gradually increased, as well as the effect of cross-flow suppression, which formed a relative balance. For all swirl chambers with variable jet nozzle angles, the thermal performance factors were all larger than 1, which indicated the heat transfer was enhanced with less friction increment.     

One—Dimensional Analysis of Thermal Choking in Case of Heat Addition in Ducts

IntroductionThe s up ersonic- combus t ion ram j et (s cramj et ) engine is one of the most prollilsing air breathing prOPulsion systems for hypersonic transports. It is essentialto the design of scramjet engines that the fueLair miX-ture remains supersonic throughout the combustor.The subsonic combustion is better than the supersonic combustion from a standpoint of efficiency, bescause total pressure losses in heated supersonic flowsare higher than those in heated subsonic flows withsame st…     

燃气轮机燃烧室与透平交互作用研究进展

当代高性能燃气轮机高温部件燃烧室和透平存在着复杂的流动传热现象.随着燃机透平进口温度不断提高,部件交互作用愈发突出,直接影响到燃烧筒和叶片材料的温度水平.总结了国内外燃机中燃烧室/透平交互作用的研究进展,包括热斑、湍流度、辐射、旋流、尾迹管理等因素,归纳了典型的理论研究、实验研究、数值模拟成果.燃烧室与透平交互作用的机理研究已取得较大进展,在真机工况下的验证与应用仍需开拓.高温部件的实验平台与计算开发是支撑设计体系建设的重要基础.     

Numerical study on the interaction characterization of rotating detonation wave and turbine rotor blades

Rotating detonation as a kind of pressure gain combustion is expected to greatly improve efficiency when applied to gas turbine engines. In this paper, the operation of rotating detonation combustor and turbine rotor blade was studied. Firstly, the analysis of the interaction between detonation wave and turbine blade shows that the compression of gas by detonation wave and reflected wave will lead to a sharp increase in the temperature at the wall of blade. When the detonation wave propagates, the oscillation amplitudes of pressure and temperature at the turbine inlet are 70% and 75% respectively, and the detonation oblique shock will change the flow trajectory of the air flow, resulting in the flow direction deviating from the incident angle. Then the comparison between detonation and deflagration shows that the total pressure of detonation is higher and will have greater work potential. The torque generated by the blades under detonation has the characteristics of high-frequency oscillation, which may be detrimental to the operation of the engine.