Enhanced Internal Heat Transfer on the Tip-Wall in a Rectangular Two-Pass Channel (AR = 1:2) by Pin-Fin Arrays

To improve gas turbine performance, the operating temperature has been increased continuously. However, the heat transferred to the turbine blade is substantially increased as the turbine inlet temperature is increased. Cooling methods are therefore much needed for the turbine blades to ensure a long durability and safe operation. The blade tip region is exposed to the hot gas flows and is difficult to cool. A common way to cool the tip is to use serpentine passages with a 180° turn under the blade tip cap taking advantage of the three-dimensional turning effect and impingement. Increasing internal convective cooling is however required to increase the blade tip life. In this article, enhanced heat transfer of a blade tip has been investigated numerically. The computational models consist of a two-pass channel with a 180° turn and arrays of pin-fins mounted on the tip-cap, and a smooth two-pass channel. Inlet Reynolds numbers range from 100,000 to 600,000. The computations are 3-D, steady, and incompressible. The detailed 3-D fluid flow and heat transfer over the tip surfaces are presented. The overall performance of the two models is evaluated. It is found that due to the combination of turning, impingement, and pin-fin crossflow the heat transfer coefficient of the pin-finned tip might be a factor of 1.84 higher than that of a smooth tip. This augmentation is achieved at the expense of a penalty of pressure drop around 35%. It is suggested that the pin-fins could be used to enhance blade tip heat transfer and cooling.     

Numerical prediction on turbine blade internal tip cooling with pin-fin and dimple/protrusion structures

ABSTRACT

Effective cooling techniques are required urgently because of high thermal loads on the blade tip region. The 180° turning bend is recognized to perform well in heat transfer on a blade tip. The thermal fluid-solid coupling models of the internal tip region with pin-fin-dimples/protrusions are established in the present paper. The local flow characteristics near the 180° turning bend, average Nu/Nu₀, and the friction loss on the impingement surfaces are obtained. The local flow field near the tip surface is influenced by the 180° turning bend, where the fluid impingement, cross-flow convection and deflection of the secondary flow exist. The average Nu of dimple/protrusion structures is increased by 3.2%-31.5% comparing to that of a smooth case. After arranging pin-fin-dimple/protrusion, the average Nu is increased to 31.2%-127.3%, much higher than dimple/protrusion structures. Furthermore, the arrangement of pin-fin-dimple/protrusion brings no significant increase in the friction, which indicates an efficient heat transfer structure with little resistance.     

Gas Turbine Blade Tip Heat Transfer and Cooling: A Literature Survey

Gas turbines are widely used for aircraft propulsion, land-base power generation, and other industrial applications like trains, marines, automobiles, etc. To satisfy the fast development of advanced gas turbines, the operating temperature must be increased to improve the thermal efficiency and output work of the gas turbine engine. However, the heat transferred to the turbine blade is substantially increased as the turbine inlet temperature is continuously increased. Thus, it is very important to cool the turbine blades for a long durability and safe operation. Cooling the blade must include cooling of the key regions being exposed to the hot gas. The blade tip region is such a critical area and is indeed difficult to cool. This results from the tip clearance gap where the complex tip leakage flow occurs and thereby local high heat loads prevail. This paper presents a literature survey of blade tip leakage flow and heat transfer, as well as research of external and internal cooling technologies. The present paper does not intend to review all published results in this field, nor review all papers from the past to now. This paper is limited to a review of recently available published works by several researchers, especially from 2001 to present, concerning blade tip leakage flow associated with heat transfer, and external or/and internal tip cooling technologies.     

Experimental Study of Heat Transfer and Pressure Loss in Channels with Miniature V Rib-Dimple Hybrid Structure

Abstract

In order to increase the thermal efficiency, the gas turbines are designed to operate at higher temperature, which requires highly efficient cooling structures for turbine blades. The dimples and ribs are effective surface structures to enhance the convective heat transfer in the gas turbine blade internal cooling. In the present study, a novel hybrid cooling structure with miniature V-shaped ribs and dimples is presented, and the heat transfer and pressure loss characteristics are obtained experimentally. The heat transfer performance of the rib–dimple structures, which include three different rib height-to-hydraulic diameter ratios of 0.017, 0.029 and 0.044 and one dimple configuration with the dimple depth-to-diameter ratio of 0.2, are studied by using the transient liquid crystal thermography technique for turbulent flow in rectangular channels within the Reynolds number range from 10,000 to 60,000. It is found that the miniature V-shaped ribs arranged upstream the dimples can significantly improve the heat transfer performance of the dimples, resulting in a more uniform heat transfer distribution on the surface. The V rib-dimple hybrid structure in the channel shows much higher heat transfer enhancement than the counterparts with only the dimples in the channels.     

基于流热耦合的燃气轮机透平叶顶冷却设计

燃气轮机透平叶顶区域存在复杂的流动和换热问题,承受很高的热负荷。为了降低透平动叶叶顶温度,在透平叶顶现有结构的基础上提出气膜冷却和气膜+内冷通道冷却两种叶顶冷却方案,并通过流热耦合计算分析冷却升级前后叶顶区域的换热和流动特性。研究发现：叶顶气膜冷却方案可有效降低叶顶温度,特别是叶顶前缘至中弦区域;而气膜＋内冷通道冷却方案基于外部气膜冷却,结合内部冷却通道设计,可进一步降低叶顶尾缘的温度;与原型叶片相比,气膜+内部冷气通道的复合冷却设计可以使叶顶尾缘最高温度降低24 K。     

The effect of start-up cycle in ceramic coating used as thermal barrier for a gas turbine bucket

The unsteady aerodynamic and aero-thermal performance of a first stage gas turbine bucket with thermal barrier coating (TBC) and internal cooling configuration were investigated by application of a three dimensional Navier–Stokes commercial turbomachinery oriented CFD-code. Convection and conduction were modeled for a super alloy blade with TBC.The CFD simulations were configured with a mesh domain including the nozzle and bucket interstage in order to accurately predict the fluid parameters at inlet and outlet of bucket. Comparisons to the gas turbine manufacturer data have permitted validation of the flow conditions at the inlet of the rotor.The effects of blade TBC surface temperature changes during a start-up cycle were simulated by means of an unsteady simulation, with unsteady inlet/outlet boundary conditions specified according to test data. The calculations include not only the fluid but also the solving of conduction within the blade, allowing for a correct modeling of the large difference of thermal inertia between the fluid and solid.The role of thermal barrier coatings (TBC) is, as their name suggests, to provide thermal insulation of the blade. A coating of about 100–400 μm can reduce the temperature by up to 200 °C. A TBC can be used either to reduce the need for blade cooling (by about 36%) increasing the turbine efficiency, while maintaining identical creep life of the substrate; or to increase considerably the creep life of the blade while maintaining level of blade cooling (and therefore allowing the blade to operate at a lower temperature for an identical turbine inlet temperature).     

燃气轮机透平叶片传热和冷却研究：内部冷却

随着燃气轮机透平进口温度的不断提高,其换热与冷却问题已然成为现代高性能燃气轮机研发中亟待解决的核心关键技术之一.透平叶片的冷却可以分为内部冷却和外部冷却,结合作者近年的研究工作,详细综述了燃气轮机透平叶片内部换热与冷却问题的研究现状与进展,着重介绍了叶片内部蛇形通道冷却、叶片内部冲击冷却和前缘的旋流冷却及尾缘柱肋冷却,指出了它们各自在相关方面需要进一步开展的工作.其中：在蛇形通道冷却方面,需要进一步研究旋转状态下蛇形通道内流动与换热特性、发展高性能的扰流装置及通道弯头结构、设计新颖高效的叶顶内部冷却结构、获得带气膜孔或冲击孔的蛇形通道内的换热与冷却特性;在叶片前缘内部冲击冷却方面,需要研究不同曲率面上的冲击冷却换热特性、旋转条件下的冲击冷却以及冲击气膜复合冷却特性;在旋流冷却方面,需要对其结构参数的影响开展进一步的广泛研究,并开展旋转状态下旋流冷却特性的研究;在尾缘柱肋冷却方面,需要进一步研究复杂流场下柱肋阵列通道中的流动换热与众敏感因子之间的关系.     

燃气轮机叶片内部扰流肋冷却方式研究进展

随着燃气透平转子进口温度的不断提高,燃气轮机叶片冷却日益重要。带有扰流肋的内部通道冷却是叶片冷却的一个重要部分。综述了内部扰流肋冷却的研究历程与研究现状,详细论述了静止状态下带肋内部通道的换热研究、旋转对带肋通道内换热的影响研究以及扰流肋与其他方式相结合的复合冷却研究。结论指出,在国内外静止状态下带肋通道内的换热研究已经很成熟,旋转对通道内流动与换热的影响是最近几年来的研究热点,而关于旋转状态下复合冷却方式的研究相对较少。优化旋转状态下内部肋结构和将内部扰流肋与其他冷却方式相结合的研究是今后的发展方向。     

舰船燃气轮机高压涡轮动叶多维热耦合设计方法

针对舰船燃气轮机复杂高效冷却叶片设计,基于压力修正算法建立冷却叶片一维管网设计方法;通过快速求解可压缩边界层微分方程获得叶片外换热边界,基于参数化的叶片网格生成方法,采用全隐式有限体积的固体导热求解方法,构建了冷却叶片的耦合传热模型,开发了耦合传热计算程序。对某高压涡轮动叶进行多维热耦合设计,确定冷却流路及冷气分布,通过三维气热耦合计算验证了设计方案的可行性,通过对比分析验证了多维热耦合设计方法对主要流通单元的流量、压力误差小于5%,具备较高的工程应用价值。     

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.     

Vortical structures and heat transfer augmentation of a cooling channel in a gas turbine blade with various arrangements of tip bleed holes

Abstract

This study investigates the internal cooling processes affected by the tip bleed holes in gas turbine blades. Double bleed holes are fixed at the center of the blade tip near the pressure side and suction side, respectively. Five different arrangements of the holes along the center line of the tip are studied. The purely double holes are set as the Baseline. The purpose of the present study is to provide a new perspective of the tip film cooling to understand the internal flow processes, vorticity evolution and the mechanism of the heat transfer augmentation. A topological analysis and the boundary layer analysis methods are introduced to better understand the tip heat transfer. The total extraction area and volume is kept at the same level for all the studied cases. The results show that the Dean vortices and the near-wall vortices induced by the secondary flow contribute to the high heat transfer coefficient on the tip surface. The mixing effect of the Dean vortices and the hole extraction helps to enhance heat transfer upstream of the tip. Different arrangement of the bleed holes can affect the internal flow processes and heat transfer performance. The suction effect of the center-line bleed hole can accelerate the near-hole flow and reduce the thickness of the boundary layer. The center-line hole fitted at the middle of the tip affects significantly the rear side of the hole. Thus, the holes aligned in the middle of the tip provide the highest heat transfer and thermal performance. The thermal performance is enhanced by up to 4.7% compared with the Baseline.     

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.     

燃气涡轮叶片内部冷却结构的设计与实验

通过某型号涡轮气冷叶片的设计与验证,建立了一种简化的气冷叶片流热解耦设计模拟方法,结合对流冷却、冲击冷却的实验结果,通过调整冷气参数,达到控制叶片表面温度的要求.在平面叶栅热风洞中,通过调整尾流板角度,控制叶片表面流动状态,满足叶栅流动的周期性条件,保证了热测量的正确性.完成了气冷叶片的设计、分析与验证的系列工作,为今后研制气冷叶片建立了扎实基础.     

不同冷气/燃气温比下高温涡轮叶片旋流冷却流固耦合特性研究

采用流固耦合方法对燃气轮机高温涡轮叶片旋流冷却结构进行数值模拟分析。探究了不同冷气/燃气温度比条件下旋流冷却的流动与传热特性、叶片前缘区域固体温度、热应力以及热应变分布。研究表明：在进气腔入口雷诺数固定的条件下,随着温度比升高,冷气密度降低,冷气流速逐渐提升,同时湍动能升高,靶面努塞尔数逐渐升高;当温度比较低时冷气的流速较低、单位时间冷气带走的热量较少,当温度比较高时冷气温度较高、单位质量冷气所能吸收的热量有限,靶面处热流密度先升高后降低。受靶面热流密度分布影响,随着温度比升高,叶片前缘固体的温度、热应力以及热应变先降低后升高。     

大型燃气轮机透平第一级静叶表面温度分布的计算方法

利用附加源项法计算叶片外表面换热系数及温度分布,编制叶片内部冷却计算程序及壁面导热程序,应用该程序计算了某大型燃气轮机第一级静叶表面温度分布,对燃气涡轮叶片先进的内外部冷却系统设计方法的消化吸收奠定了基础。     

透平叶顶气膜冷却效果数值研究

采用三维数值模拟方法,研究了GE E3发动机第一级透平动叶叶顶间隙内的气膜流动与换热特性,评估了气膜吹风比M分别为0.5、1.0和1.5时,对叶顶换热系数以及冷却效率的影响.计算结果表明:叶顶气膜冷却空气改变了叶顶泄漏流动特性,随着吹风比的增加,叶顶间隙内的泄漏流动区域不断缩小,从而导致叶顶间隙泄漏量不断减小;随着气膜冷却吹风比的增大,叶顶平均换热系数逐步降低;在M=1时,冷却效果最佳.     

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.     

带有冷气掺混的三级透平气动性能数值研究

借助NUMECA数值仿真软件,以某型燃气轮机的三级透平作为计算模型,对其在冷却气体掺混前后的流场进行了数值模拟。考虑到工质物性的影响,采用了变比热高温燃气作为计算工质。同时,针对燃气轮机透平进口的变工况问题,选取不同的透平进口总压值进行数值计算。结果表明,冷却气体的加入使得级损失增大,每列叶片流道出口速度或相对速度减小,下游叶片进口气流角减小;在三级透平冷气掺混时改变进口总压值,每列叶片流道的进口气流角几乎不变,除第三级动叶的激波损失与尾迹损失增大外,其余叶片流道的能量损失变化不明显。     

Experimental Studies on Heat Transfer in the Tip Gap of a Sectorial Turbine Cascade

ＥｘｐｅｒｉｍｅｎｔａｌＳｔｕｄｉｅｓｏｎＨｅａｔＴｒａｎｓｆｅｒｉｎｔｈｅＴｉｐＧａｐｏｆａＳｅｃｔｏｒｉａｌＴｕｒｂｉｎｅＣａｓｃａｄｅＥｘｐｅｒｉｍｅｎｔａｌＳｔｕｄｉｅｓｏｎＨｅａｔＴｒａｎｓｆｅｒｉｎｔｈｅＴｉｐＧａｐｏｆａＳｅｃｔｏｒｉａ...     

Heat transfer in internal channel of a blade: Effects of rotation in a trailing edge cooling system

The aerothermal performance of a trailing edge (TE) internal cooling system of a high pressure gas turbine blade was evaluated under stationary and rotating conditions.The investigated geometry consists of a 30:1 scaled model reproducing a typical wedge shaped discharge duct with one row of enlarged pedestals.The airflow pattern inside the device simulates a highly loaded rotor blade cooling scheme with a 90 [deg] turning flow from the radial hub inlet to the tangential TE outlet.Two different tip configurations were tested,the first one with a completely closed section,the second one with a 5 holes outlet surfaces discharging at ambient pressure.In order to assess rotation effects,a rotating test rig,composed of a rotating arm holding both the PMMA TE model and the instru-mentation,was purposely developed and manufactured.A thin Inconel heating foil and wide band Thermo-chromic Liquid Crystals are used to perform steady state heat transfer measurements on the blade pres-sure side.A rotary joint ensures the pneumatic connection between the blower and the rotating apparatus;more-over several slip rings are used for both instrumentation power supply and thermocouple connection.A parallel CFD analysis involving steady-state RANS modeling was conducted to allow an insight of the flow field inside the redirecting channel and the interpedestal ducts to better interpret the developing vortical structures.Low-Reynolds grid clustering permits to integrate up to the wall both the momentum and the thermal boundary layer.Calculations were performed by means of an in-house developed pressure based solver exploiting the k-ω SST turbulence model implemented in the framework of the open-source finite volume discretization toolbox OpenFOAM?.Analyzed flow conditions correspond to Reynolds number of 20000 in the hub inlet section and angular speed varies to obtain rotation numbers in the range from 0 to 0.3.The orientation of the rotation axis is orthogonal to the heated surface as to resemble a 90 [deg] blade metal angle.Results are reported in terms of de-tailed heat transfer coefficient 2D maps on the suction side surface as well as spanwise profiles inside the pedestal ducts.