The interior heat transfer characteristics of gas turbine blade due to sparse film cooling holes

The film cooling technique is one of the most useful cooling methods. At present, the midchord region of gas turbine blades in an aeroengine often adopt a sparse film cooling technique and impingement cooling technique at the same time. So the interior heat transfer characteristics on the inner side of blades due to the sparse film cooling holes have become a very complicated and interesting problem. In this paper, the heat transfer characteristics of impingement‐cooling have been investigated experimentally. Through lots of experimental data, the effect of flow parameters and geometric parameters on heat transfer characteristics has been studied. Correlation equations obtained show good agreement with experimental data. © 2005 Wiley Periodicals, Inc. Heat Trans Asian Res, 34(3): 197–207, 2005; Published online in Wiley InterScience ( www.interscience.wiley.com ). DOI 10.1002/htj.20052     

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.     

Flow structure and heat exchange analysis in internal cooling channel of gas turbine blade

This paper presents the study of the flow structure and heat transfer, and also their correlations on the four walls of a radial cooling passage model of a gas turbine blade. The investigations focus on heat transfer and aerodynamic measurements in the channel, which is an accurate representation of the configuration used in aeroengines. Correlations for the heat transfer coefficient and the pressure drop used in the design of radial cooling passages are often developed from simplified models. It is important to note that real engine passages do not have perfect rectangular cross sections, but include corner fillet, ribs with fillet radii and special orientation. Therefore, this work provides detailed fluid flow and heat transfer data for a model of radial cooling geometry which possesses very realistic features.     

Influence of internal channel geometry of gas turbine blade on flow structure and heat transfer

This paper presents the study of the influence of channel geometry on the flow structure and heat transfer,and also their correlations on all the walls of a radial cooling passage model of a gas turbine blade.The investigations focus on the heat transfer and aerodynamic measurements in the channel,which is an accurate representation of the configuration used in aeroengines.Correlations for the heat transfer coefficient and the pressure drop used in the design of internal cooling passages are often developed from simplified models.It is important to note that real engine passages do not have perfect rectangular cross sections,but include a comer fillets,ribs with fillet radii and a special orientation.Therefore,this work provides detailed fluid flow and heat transfer data for a model of radial cooling geometry which has very realistic features.     

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.     

Film cooling on a gas turbine blade pressure side or suction side with axial shaped holes

The film cooling effectiveness on the surface of a high pressure turbine blade is measured using the pressure sensitive paint (PSP) technique. Four rows of axial laid-back, fan-shaped cooling holes are distributed on the pressure side while two such rows are provided on the suction side. The coolant is only injected to either the pressure side or suction side of the blade at five average blowing ratios ranging from 0.4 to 1.5. The presence of wakes due to upstream vanes is simulated by placing a periodic set of rods upstream of the test blade. Effect of the upstream wakes is recorded at four different phase locations with equal intervals along the pitch-wise direction. The freestream Mach numbers at cascade inlet and exit are 0.27 and 0.44, respectively. Results reveal that the tip leakage vortices and endwall vortices sweep the coolant film on the suction side to the midspan region. The film cooling effectiveness on the suction side is usually higher than that on the pressure side except the regions affected by the secondary vortices. The presence of upstream wakes results in lower film cooling effectiveness on the blade surface. The moderate blowing ratios (M = 0.6 or M = 0.9) give higher film cooling effectiveness immediately downstream of the film cooling holes. Further downstream of the holes, higher blowing ratios cover wider surface area.     

Effects of rim location, rim height, and tip clearance on the tip and near tip region heat transfer of a gas turbine blade

Effects of the rim height and the tip gap clearance on the heat transfer coefficients on the blade tip and near tip regions were measured with two different rim geometries. The heat transfer coefficient distributions were measured using the transient single color capturing liquid crystals technique. Rims were located along (a) the pressure and the suction side (full-rim case) and (b) the suction side of the blade tip (suction side rim case). The rim heights were (a) 2.1%, (b) 4.2%, and (c) 6.3% and the blade tip gap clearances were (a) 1.0%, (b) 1.5%, and (c) 2.5% of the blade span. Tests were performed on a five-bladed linear cascade placed in a blowdown facility. The overall pressure ratio, inlet total pressure to exit static pressure, was 1.2, and the Reynolds number based on the exit velocity and the axial cord length was 1.1 × 10⁶. The turbulence intensity level at the cascade inlet was 9.7%, and the inlet and exit Mach number were 0.25 and 0.59, respectively. It was found that higher rims reduce the heat transfer coefficients on the tip and shroud, but the reduction on the pressure and suction sides was not significant. The suction side rim case provided lower heat transfer coefficients on the blade tip and near tip regions than the full-rim case.     

Effect of various rib arrangements on heat transfer in a semicylinder channel with effusion flow

The flow and heat transfer characteristics of various rib configurations on a concave channel surface with effusion holes were investigated. A semicylindrical channel with three rows of effusion holes was used to simplify the blade leading edge and eight kinds of ribs were attached on the internal concave surface for comparison. Continuous and broken ribs were both applied at 90°, as were upstream-pointed V-shaped and downstream-pointed V-shaped ribs. The Reynolds-averaged Navier–Stokes equation was solved using commercial software. The result included the divided-area-averaged and local Nusselt number distribution; the overall average Nusselt number on the concave surface is also discussed.     

Numerical simulation of steady state heat transfer in a ceramic-coated gas turbine blade

As gas turbine entry temperature (TET) increases, thermal loading on first stage blades increases and, therefore, a variety of cooling techniques and thermal barrier coatings (TBCs) are used. In the present work, steady state blade heat transfer mechanisms were studied via numerical simulations. Convection and radiation to the blade external surface were modeled for a super alloy blade with and without a TBC. The effects of surface emissivity changes, partial TBC coatings and uncertainties in external heat transfer coefficient were also simulated. The results show that at 1500 K TET, radiation heat transfer rate from gas to an uncoated blade is 8.4% of total heat transfer rate which decreases to 3.4% in the presence of a TBC. The TBC blocks radiation, suppresses metal temperatures and reduces heat loss to the coolant. These effects are more pronounced at higher TETs. With selective coating, substantial local temperature suppression occurs. In the presence of radiation and/or TBC, the uncertainties in convection heat transfer coefficient do not have a significant effect on metal temperatures.     

Effect of rotation on leading edge region film cooling of a gas turbine blade with three rows of film cooling holes

Effect of rotation on detailed film cooling effectiveness distributions in the leading edge region of a gas turbine blade with three showerhead rows of radial-angle holes were measured using the Pressure Sensitive Paint (PSP) technique. Tests were conducted on the first-stage rotor blade of a three-stage axial turbine at three rotational speeds. The effect of the blowing ratio was also studied. The Reynolds number based on the axial chord length and the exit velocity was 200,000 and the total to exit pressure ratio was 1.12 for the first-stage rotor blade. The corresponding rotor blade inlet and exit Mach number was 0.1 and 0.3, respectively. The film cooling effectiveness distributions were presented along with the discussions on the influences of rotational speed, blowing ratio, and vortices around the leading edge region. Results showed that different rotation speeds significantly change the film cooling traces with the average film cooling effectiveness in the leading edge region increasing with blowing ratio.     

Fluid flow characteristics and convective heat transfer improvement on a gas turbine blade

The flow field features and heat transfer enhancement are investigated on a gas turbine blade by applying the jet impingement cooling method. The distribution of the flow field and the Nusselt number (Nu) was determined on the targeted surface in the cooling channel. The injection holes of different shapes, such as circular, square, and rectangular were considered. The Reynolds numbers (Re) of the airflow in the range of 2000–5000 and aspect ratios of 0.5–2 were particularly focused. The flow vortices and recirculation in the cooling channel and their influence on the heat transfer enhancement were analyzed in detail under different airflow and geometric conditions. Decreasing the ratio of the distance between jet-to-target plate to the diameter of the jet orifice (H/d) increased the heat transfer rate and produced high-intensity vortices and recirculation zones. It was noticed that the formation and generation of vortices and recirculation have important effects on the convective heat transfer rate at the impingement surface. Local Nusselt number, formation of complex vortices, and airflow recirculation in the cooling channel decreased with the increase in the distance between the jet hole and the targeted surface. It was found that with the increase in the Reynolds number of the jet, heat transfer between cold airflow and the targeted surface increased. Moreover, it was observed that the cooling performance of the round and square jet holes was better than the rectangular holes.     

某燃气轮机涡轮叶片复合冷却结构优化设计

为了提高涡轮叶片设计效率,搭建了基于一维管网理论的iSIGHT优化设计平台,获得了叶片的优化结构,并对其进行三维仿真计算分析。研究表明：优化前后叶片的气动性能变化不大,总压损失仅在端壁处略有差别;优化后涡轮叶片的壁面温度分布更加均匀,壁面的最高温度降低,温度梯度减小,最大相对温差降低10%左右;在降低叶片热应力的同时相对冷却效率提高1.0%。     

Conjugate heat transfer investigation on the cooling performance of air cooled turbine blade with thermal barrier coating

A hot wind tunnel of annular cascade test rig is established for measuring temperature distribution on a real gas turbine blade surface with infrared camera. Besides, conjugate heat transfer numerical simulation is performed to obtain cooling efficiency distribution on both blade substrate surface and coating surface for comparison. The effect of thermal barrier coating on the overall cooling performance for blades is compared under varied mass flow rate of coolant, and spatial difference is also discussed. Results indicate that the cooling efficiency in the leading edge and trailing edge areas of the blade is the lowest. The cooling performance is not only influenced by the internal cooling structures layout inside the blade but also by the flow condition of the mainstream in the external cascade path. Thermal barrier effects of the coating vary at different regions of the blade surface, where higher internal cooling performance exists, more effective the thermal barrier will be, which means the thermal protection effect of coatings is remarkable in these regions. At the designed mass flow ratio condition, the cooling efficiency on the pressure side varies by 0.13 for the coating surface and substrate surface, while this value is 0.09 on the suction side.     

Film cooling performance of converging slot-hole rows on a gas turbine blade

Experimental tests have been performed to investigate the film cooling performance of converging slot-hole (console) rows on the turbine blade. Film cooling effectiveness of each single hole row is measured under three momentum flux ratios based on the wide-band liquid crystal technique. Measurements of the cooling effectiveness with all the hole rows open are also carried out under two coolant–mainstream flux ratios. Film cooling effectiveness of cylindrical hole rows on the same blade model is measured as a comparison. The results reveal that the trace of jets from both consoles and cylindrical holes is converging on the suction surface and expanding on the pressure surface by the influence of the passage vortex, while the influence of passage vortex on the jets from consoles is weaker. The film coverage area and the film cooling effectiveness of single/multiple console row(s) are much larger than those of single/multiple cylindrical hole row(s). When the console row is discrete and the diffusion angle of the console is not very large, the adjacent jets cannot connect immediately after ejecting out of the holes and the cooling effectiveness in the region between adjacent holes is relatively lower. On the pressure surface, the film cooling effectiveness of console rows increases notably with the increasing of momentum flux ratio or coolant–mainstream flux ratio. But on the suction side, the increase in cooling effectiveness is not very notable for console row film cooling as the coolant flux increases. Moreover, for the film cooling of single console row at the gill region of the suction surface, the jets could lift off from the blade surface because of the convex geometry of the suction surface.     

Visualization of heat transfer distributions in a gas turbine scroll

To visualize heat transfer distributions in systems with complex internal geometries, an experimental technique using a combination of the transient method and the hysteresis effect of thermopaint was developed. A mercury compound based thermopaint was used as a temperature indicator. Features of the paint are its reusability and its hysteresis nature. Isothermal lines visualized by the thermopaint are preserved after the experiment by utilizing the hysteresis nature. Also heat transfer on those parts that are hidden behind other parts can be visualized. Effects of initial and air temperature on the measurement uncertainty were evaluated. With this method, local heat transfer coefficients were obtained on the model scroll of a single can type combustor. © 1998 Scripta Technica, Heat Trans Jpn Res, 27(3): 229–242, 1998     

车用PEMFC直流道散热特性研究

质子交换膜燃料电池(PEMFC)的散热对其性能有很大影响。文章利用Gambit软件建立带冷却通道的PEMFC模型,使用计算流体力学软件Fluent中的PEM模块进行数值模拟计算。通过改变冷却通道进口处冷却水的流速和温度,对质子交换膜内温度和冷却水出口处温度进行了分析。数据表明,冷却水的流速和温度对PEM内温度分布都有一定影响。为使PEMFC正常稳定工作,冷却水流速不宜过小、温度不宜过低。     

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

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Endwall heat transfer and film cooling measurements in a turbine cascade with injection upstream of leading edge

Detailed heat transfer measurements were conducted on the endwall surface of a large‐scale low‐speed turbine cascade with single and double row injection on the endwall upstream of leading edge. Local film cooling effectiveness and the heat transfer coefficient with coolant injection were determined at blowing ratios 1.0, 2.0, and 3.0. In conjunction with the previously measured flow field data, the behaviors of endwall film cooling and heat transfer were studied. The results show that endwall film cooling is influenced to a great extent by the secondary flow and the coverage of coolant on the endwall is mainly determined by the blowing ratio. An uncovered triangle‐shaped area with low effectiveness close to pressure side could be observed at a low blowing ratio injection. The averaged effectiveness increases significantly when injecting at medium and high blowing ratios, and uniform coverage of coolant on the endwall could be achieved. The averaged effectiveness could be doubled in the case of double row injection. It was also observed that coolant injection made the overall averaged heat transfer coefficient increase remarkably with blowing ratio. It was proven that film cooling could reduce endwall heat flux markedly. The results illustrate the need to take such facts into account in the design process as the three‐dimensional flow patterns in the vicinity of the endwall, the interactions between the secondary flow and coolant, and the augmentation of heat transfer rate in the case of endwall injection. © 2004 Wiley Periodicals, Inc. Heat Trans Asian Res, 33(3): 141–152, 2004; Published online in Wiley InterScience ( www.interscience.wiley.com ). DOI 10.1002/htj.20007     

Heat transfer measurement techniques of film cooling in gas turbine blades

An overview of the experimental techniques employed or developed for the measurement of local and mean heat transfer coefficients and adiabatic wall effectiveness from film cooled surfaces is presented. The scope of this work is confined to heat transfer techniques applied to film cooling of gas turbine blades, steady state and transient. The latter technique have significant advantages over the former in that it yields results at parameters duplicating those at the full-scale operating engine conditions, although the former technique offers simplicity.     

Mass transfer mesurement techniques of film cooling in gas turbine blades

This paper presents an overview of the analogous mass transfer experimental techniques that are used or developed for the measurement of local and mean heat-mass transfer coefficients and adiabatic wall effectiveness, chiefly for the film-cooled blades of gas turbines. The mass transfer techniques yield comprehensive results that are more difficult to obtain using the conventional heat transfer methods. They also provide results such that the errors associated with direct heat transfer methods are eliminated.