Leading Edge Region Research Articles

Swirl Impinging Cooling on an Airfoil Leading Edge Model at Large Reynolds Number

Jet impingement cooling has been extensively investigated due to its significant applications on the airfoil leading edge region; however, most of which are about normal jet impingement. The systematic research on swirl jet impinging cooling on leading edge is relatively rare. This study comprehensively investigated the heat transfer distribution of swirl jet impingement with one row of tangential jets. The location of the cross-over jets is offset from the centerline toward either suction or pressure side. Five jet Reynolds numbers varying from 10,000 to 80,000 are tested to reach real engine cooling condition. Jet plates with jet-to-jet spacing (s/d = 2, 4, and 8) and the ratio of surface diameter-to-jet diameter (D/d = 4, 6.6, and 13.3) are tested. We conducted the experiments with a test matrix of 45 cases. The optimum geometric parameters of the jet plate are revealed. Results indicate that for a given Reynolds number, the jet plate configuration with D/d = 4 and s/d = 2 provides the highest Nusselt number profile than the other jet plate configurations, while the jet plate configuration with D/d = 13.3 and s/d = 8 provides the lowest Nusselt number profiles. The best heat transfer region shifts by varying the jet plate configuration depending on the strength of swirl flow. Additionally, correlation of tangential jet impingement has been developed to predict the area-averaged Nusselt number, which is useful for airfoil leading edge cooling design and heat transfer analysis.

An experimental investigation on cooling characteristics of a vane laminated end-wall with axial-row layout of film-holes

A laminated cooling structure with axial-row layout of film-holes was employed in the design of a first-stage turbine vane end-wall. Different internal elements were placed in the regions near leading edge (LE), pressure side (PS) and suction side (SS). Conjugate heat transfer effects of fluid and solid at the laminated cooling end-wall (LCE) were measured under low and engine-near temperature ratios (TRs) of mainstream-to-coolant, respectively. The detailed comparison of overall cooling effectiveness and surface temperature gradients with a traditional film cooling end-wall (FCE) at three coolant-to-mainstream mass flow ratios (MFRs) and two TRs revealed that the LCE cannot only decease the temperature of the hot surface and the area of the regions difficult to be cooled by simple film cooling, but also decrease the temperature gradient within the end-wall, and improve the reliability and durability of the end-wall structure thereby. The overall cooling effectiveness of LCE can be increased by 15% in LE and SS regions and about 25% in PS region, when MFR is larger than 0.8%. The temperature gradients over the entire LCE can be reduced by about 20% at MFR = 1.2%. In addition, at the engine-near TR condition, these benefits of LCE mentioned above are still remarkable.

Leading-Edge Flow Sensing for Aerodynamic Parameter Estimation

The identification of inflow air-data quantities such as the airspeed, angle of attack, and local lift coefficient on various sections of a wing or rotor blade is beneficial for load monitoring, aerodynamic diagnostics, and control on devices ranging from air vehicles to wind turbines. Real-time measurement of aerodynamic parameters during flight provides the ability to enhance aircraft operating capabilities while preventing dangerous stall situations. This Paper presents a novel leading-edge flow sensing algorithm for the determination of the air-data parameters using discrete surface pressures measured at a few ports in the vicinity of the leading edge of a wing or blade section. The approach approximates the leading-edge region of the airfoil as a parabola and uses pressure distribution from the exact potential-flow solution for the parabola to fit the pressures measured from the ports. Pressures sensed at five discrete locations near the leading edge of an airfoil are given as input to the algorithm to solve for the model parabola-flow problem using a simple nonlinear regression. The algorithm directly computes the inflow velocity and the stagnation-point location. A look-up table approach is adopted to deduce the section angle of attack and lift coefficient. The performance of the algorithm is assessed using experimental data in the literature for airfoils under different flow conditions. The results show good correlation between the actual and predicted aerodynamic quantities within the prestall regime, even for rotating blade sections.

Numerical evaluation of heat/mass transfer analogy for leading edge showerhead film cooling on a first-stage vane

A numerical study was conducted to evaluate the applicability of the heat/mass transfer analogy to leading edge film cooling on a first-stage vane. A typical showerhead configuration was taken into account, thus including four staggered rows of evenly spaced cylindrical holes, angled at 45° towards the tip. The investigation was carried out at low speed (isentropic exit Mach number of Ma2is = 0.2), inlet turbulence intensity of Tu1 = 1.6% and 13%, and three different blowing ratios (BR = 2.0, 3.0 and 4.0). Steady-state simulations were conceived to mimic the pressure sensitive paint (PSP) technique aiming to measure the adiabatic film cooling effectiveness (η) in the leading edge region, under a density ratio of DR = 1 and 1.5. With the mainstream flow assumed to be air at ambient temperature, the desired DR was obtained by injecting coolant as air at different temperature from the mainstream (namely, heat transfer approach) or as isothermal foreign gas (namely, mass transfer approach). Despite the dominant effect of the DR on η, results indicated that the heat/mass transfer properties of the coolant are to be considered when comparing film cooling performance, at matched BR conditions. At low inlet turbulence intensity, the heat/mass transfer analogy was found to be applicable only to the lowest BR of 2.0. Its validity was extended to the medium BR of 3.0, in case of DR = 1.5, provided that the leading edge is approached by high turbulence intensity flow.

Effect of leading-edge geometry on the aerodynamics and heat transfer in the stagnation region of uncooled turbine blades

The flow in the leading-edge stagnation region of uncooled turbine aerofoils is studied. The stagnation region is modelled based on the Hiemenz flow solution, following Holley and Langston (J. Turbomach. 130: 021001, 2008). These results are applied to the JT9D turbine blade and confirmed by computations. It is also shown that the heat transfer at the stagnation point is bounded by the Hiemenz flow and plane stagnation-point potential flow heat transfer solutions. The computations are further extended for a range of Reynolds numbers and freestream turbulence intensities. It is seen that for a wide range of these parameters considered in the present study and data collected from published literature, a majority of the data points are within these bounds. The leading-edge geometry of the JT9D turbine blade is modified with elliptical geometries. Blunter-leading-edge geometries have lower values of heat transfer at the stagnation region; however, their blade profile loss coefficients are higher. The off-design performance characteristics of the turbine blades are also computed. The results presented in this paper will be useful in designing leading-edge geometries for reduced heat transfer in the stagnation region of uncooled turbine blades.

Influence of fillet shapes on secondary flow field in a transonic axial flow turbine stage

Numerical experiments were conducted to investigate the effect of leading edge modifications via fillet shapes near vane/blade-endwall juncture in a transonic environment within the highly loaded high pressure turbine stage. The investigated fillet shapes were designed based on geometric parameters: leading edge radius and included angle. The geometrical modifications were achieved to achieve variation in fillet radii at vane/blade endwall juncture along the stream-wise direction, namely variable fillet and constant fillet. Further their influences were studied in both nozzle guide vane and rotor passage secondary flow field. Computational Fluid Dynamics (CFD) method was used to resolve the flow features inside the turbine passage for planar and fillet cases. The presented data highlight the secondary flow features and their behavior using topological properties of flow field aided with the streamline and iso-contour plots. The flow-field results show a significant reduction in the total pressure losses associated with the horse shoe vortex near the leading edge region as the fillet radii are varied. Overall in both vane and rotor passages, variable fillet outperforms the constant fillet by reducing the penetration length of three-dimensional regimes along its span, mitigating the boundary layer growth and improving the loss coefficients.

Experimental investigation on the effects of rotation and the blowing ratio on the leading-edge film cooling of a twist turbine blade

An experimental investigation has been performed to investigate the effects of the rotation and blowing ratio on the film cooling effectiveness distributions of the leading-edge regions of a twist gas turbine blade using a thermochromic liquid crystal (TLC) technique. The experiments were carried out at three rotating speeds, including 400 rpm (positive incidence angle), 550 rpm (zero incidence angle), and 700 rpm (negative incidence angle). The averaged blowing ratio ranged from 0.5 to 2.0. CO2 was used as the coolant to ensure that the coolant-to-mainstream ratio was equal to 1.56. The Reynolds number, based on the mainstream velocity of the turbine outlet and the rotor blade chord length, was 6.08 × 104. The effects of the rotating speed and the blowing ratio were analyzed based on the film cooling effectiveness distribution. The results show that rotating speed plays an indispensable role in determining the film cooling effectiveness of distributions on the leading edge. The position of the stagnation line moves from the pressure side (PS) to the suction side (SS) via an increase in rotating speed. Under the same blowing ratio, the area-averaged film cooling effectiveness increases monotonously with an increase in rotating speed. Under the same rotating speed, the area-averaged film cooling effectiveness increases with the increase in blowing ratio. More details about the effects of the rotation speed and blowing ratio on the spanwise averaged film cooling effectiveness of the leading-edge region are shown in this study.

Flow analysis of materials in friction stir welding

AbstractThe aim of this study was to analyze the flow behavior of materials in friction stir welding (FSW) of magnesium (Mg) alloy to aluminum (Al) alloy materials at a low rotational speed of 300 rpm and at a low welding speed of 50 mm/min. It was observed that during the dissimilar welding at 300 rpm, some of the material of the advancing side (AS) rotated almost 360° and transferred to the position just behind the previous position, i.e. initial AS position and some of the material of leading edge region transferred to the trailing edge region and such a change continued. Material transfer takes place layer by layer from front to back portions of the FSW tool; a thin layer of material adhered to the tool in the front portion for every half rotation and transferred to the back portion of the tool and this repeated in subsequent rotations. Wrinkle formation began at the front side of the forward moving tool and as the material transferred to backside they became an onion ring pattern, and this was inevitable because the formation of onion rings instantly in the backside region of the tool was clearly impossible.

Heat Transfer Augmentation and Cooling of a Turbine Blade Using an Innovative Converging-Diverging Ducts – A CFD Study

Gas turbines are used in power plants to supply peak loads in hydro-plants and they are also used in jet, aircraft and ships for propulsion. Turbine blades extract energy from high temperature and pressure of combustion gases. In order to endure this extreme condition of temperature, gas turbine blades are often cooled internally. This ensures that the entry temperature of the combustion gases could be increased which in turn increases the thermal efficiency of a gas turbine blade and the turbine could work for a reasonable lifetime. With this desire to increase entry temperature, converging-diverging (CD) ducts are incorporated in the leading edge region as cooling passages for analysis. The CD ducts are used so that fluid passes through several fluid passages which increase its velocity and ultimately provides enhanced cooling of the turbine blades. From the analysis, it is observed that ducts with minor diameter of 7mm, major diameter of 8.5mm, and pitch of 5mm result in better cooling of turbine blades and lead to optimum thermal performance. Moreover, this study also reveals that incorporating turbulators has a significant impact on cooling of turbine blades. Turbulator arrangement with relative roughness (p/e) ratio of 2 for pitch length (ph) of 5mm gives enhanced cooling and it reduces blade wall temperature by 17%. Copyright © 2018 Praise Worthy Prize S.r.l. - All rights reserved.

Experimental investigations of the effects of the injection angle and blowing ratio on the leading-edge film cooling of a rotating twisted turbine blade

Experimental investigations were performed to study the effects of the injection angle of cylindrical holes and the blowing ratio on the leading-edge-region film cooling of a twisted turbine blade under rotating conditions. The experiments were carried out at a test facility with a 1-stage turbine using the thermochromic liquid crystal (TLC) technique. All experiments were performed at a rotating speed of 574 rpm with an average blowing ratio ranging from 0.5 to 2.0. The Reynolds number was fixed at 6.3378 × 104 based on the mainstream velocity of the turbine outlet and the rotor blade chord length. CO2 was used as the coolant to achieve a coolant-to-mainstream density ratio of 1.56. The film-hole injection angles tested were 30°, 45° and 60°. The results show that both the injection angle and the blowing ratio have significant impacts on film cooling effectiveness. For α = 30° and α = 45°, the radial average film cooling effectiveness increases as the blowing ratio increases in all regions. For α = 60°, this effectiveness first increases and then decreases as the blowing ratio increases, with the case of M = 1.5 yielding the best average cooling performance. At each blowing ratio, the α = 30° case always yields the highest streamwise average film cooling effectiveness in the region of −4.3 < X/D < 2. For 2.75 < X/D < 3.75, the effectiveness first increases and then decreases as the injection angle increases. For α = 30° and α = 45°, the area average film cooling effectiveness monotonously increases as the blowing ratio increases. For α = 60°, this effectiveness first increases and then decreases as the blowing ratio increases from 0.5 to 2.0, with the best blowing ratio M = 1.5. Under the same blowing ratio, the α = 30° case always yields the highest area average film cooling effectiveness in the leading edge region.

Effects of unsteady wakes on heat transfer of blade tip and shroud

An experimental study has been conducted to investigate the heat-transfer characteristics of blade tips and shrouds with and without unsteady wakes. Depending on the presence of unsteady wakes, the local heat/mass-transfer coefficients of the tip and shroud were measured using the naphthalene sublimation method. Wakes from unsteady blades were modeled as wakes generated from moving cylindrical rod bundles. Test conditions were set to a Reynolds number of 100,000, based on an inlet velocity of 11.4 m/s and the axial chord length. The Strouhal number was varied from 0 to 0.22. For the case without unsteady wakes (St = 0), high heat/mass-transfer coefficients appeared in regions where various flow patterns, such as flow reattachment, swirling flow, and vortices, occurred. Unsteady wakes (St = 0.22) made high turbulence intensity of the tip leakage flow, and flow patterns ranging from flow reattachment to tip leakage vortex in the tip and shroud were changed and dispersed in the presence of unsteady wakes, thus changing the heat/mass-transfer distributions in these areas. Due to the high thermal load on the tip and shroud under unsteady wake conditions, more detailed cooling designs must be considered, ranging from film cooling holes to the installation of additional structures on the blade tip, especially in the leading edge region and mid-region I of the tip, and the mid-region I and trailing edge region of the shroud.

Hybrid RANS/LES Simulation of Corner Stall in a Linear Compressor Cascade

Current design-cycle Reynolds-averaged Navier–Stokes (RANS) based computational fluid dynamics (CFD) methods have the tendency to over-predict corner-stall events for axial-flow compressors operating at off-design conditions. This shortcoming has been demonstrated even in simple single-row cascade configurations. Here we report on the application of hybrid RANS/large eddy simulation (LES), or detached eddy simulation (DES), for simulating the corner-stall data from the linear compressor cascade work conducted at Ecole Centrale de Lyon. This benchmark data set provides detailed loss information while also revealing a bimodal behavior of the separation which, not surprisingly, is also not well modeled by RANS. The hybrid RANS/LES results presented here predict bimodal behavior similar to the data only when special treatment is adopted to resolve the leading-edge endwall region where the horseshoe vortex (HV) forms. The (HV) is shown to be unstable, which produces the bimodal instability. The DES simulation without special treatment or refinement in the HV region fails to predict the bimodal instability, and thus the bimodal behavior of the separation. This, in turn, causes a gross over-prediction in the scale of the corner-stall. The HV region is found to be unstable with rolling of the tertiary vortex (TV) over the secondary vortex and merging with the primary HV. With these flow dynamics realized in the DES simulations, the corner stall characteristics are found to be in better agreement with the experimental data, as compared to RANS and standard DES approaches.

Influence of Turbine Blade Leading Edge Profile on Film Cooling With Shaped Holes

This paper presents the turbine blade leading edge model film cooling effectiveness with shaped holes, using the pressure sensitive paint (PSP) mass transfer analogy method. The effects of leading edge profile, coolant to mainstream density ratio, and blowing ratio are studied. Computational simulations are performed using the realizable k–ɛ (RKE) turbulence model. Effectiveness obtained by computational fluid dynamics (CFD) simulations is compared with experiments. Three leading edge profiles, including one semicylinder and two semi-elliptical cylinders with an after body, are investigated. The ratios of major to minor axis of two semi-elliptical cylinders are 1.5 and 2.0, respectively. The leading edge has three rows of shaped holes. For the semicylinder model, shaped holes are located at 0 deg (stagnation line) and ±30 deg. Row spacing between cooling holes and the distance between impingement plate and stagnation line are the same for three leading edge models. The coolant to mainstream density ratio varies from 1.0 to 1.5 and 2.0, and the blowing ratio varies from 0.5 to 1.0 and 1.5. Mainstream Reynolds number is about 100,000 based on the diameter of the leading edge cylinder, and the mainstream turbulence intensity is about 7%. The results provide an understanding of the effects of leading edge profile on turbine blade leading edge region film cooling with shaped hole designs.

Experimental Investigation on Blunt-Edged UTM Delta Wing VFE-2 Configurations at Low Reynolds Number

The flow behaviour over the upper surface of a blunt-edged delta wing is mainly governed by a complicated leading edge vortex. This paper presents an experimental investigation on blunt-edged UTM VFE-2 delta wing model at low Reynolds numbers. The primary vortex for sharp-edged delta wing develops in the apex region even at low angles of attack. However, this situation does not happen for blunt-edged wings; the primary vortex in the leading edge region develops at a certain chordwise position on the wing. The primary vortex progresses upstream or downstream depending on the leading edge profile, the angle of attack and Reynolds number. A decrease in Reynolds number will accelerate the upstream progression of the primary vortex towards the wing apex. In the VFE-2 research group, many experiments were conducted at Reynolds number of 1×10^6 and above. Thus, the main objective of this study was to investigate the upstream progression of leading edge primary vortex on a blunt-edged delta wing at Reynolds numbers of 0.5×10^6, 0.75×10^6 and 1.0×10^6. Wind tunnel experiments were performed at three different flow velocities, 8.7 m/s, 13.1 m/s and 17.5 m/s corresponding to Reynolds numbers of 0.5×10^6, 0.75×10^6 and 1.0×10^6. Two measurement techniques were employed on the upper surface of the wing, i.e. experimental surface pressure and tuft techniques. Experiments were also conducted on two different leading edge profiles namely medium and large radius profiles with different leading edge radius-to-wing chord ratio. Pressure coefficients on the upper surface of the wing were plotted to observe the characteristic primary vortex, the upstream progression of the primary vortex and also the characteristics of vortex breakdown. The results obtained indicate that the primary vortex rapidly formed when the Reynolds number was decreased. Analysis in the apex region showed that the attached flow still existed even at a very low Reynolds number of 0.5 × 10^6.

Turbine Blade Leading Edge Cooling With One Row of Normal or Tangential Impinging Jets

Jet impingement cooling has been extensively used in the leading edge region of a gas turbine blade. This study focuses on the effect of jet impinging position on leading edge heat transfer. The test model is composed of a semicylindrical target plate, side exit slots, and an impingement jet plate. A row of cylindrical injection holes is located along the axis (normal jet) or the edge (tangential jet) of the semicylinder, on the jet plate. The jet-to-target-plate distance to jet diameter ratio (z/d) is 5 and the ratio of jet-to-jet spacing to jet diameter (s/d) is 4. The jet Reynolds number is varied from 10,000 to 30,000. Detailed impingement heat transfer coefficient distributions were experimentally measured by using the transient liquid crystal (TLC) technique. To understand the thermal flow physics, numerical simulations were performed using Reynolds-averaged Navier–Stokes (RANS) with two turbulence models: realizable k–ε (RKE) and shear stress transport k–ω model (SST). Comparisons between the experimental and the numerical results are presented. The results indicate that the local Nusselt numbers on the test surface increase with the increasing jet Reynolds number. The tangential jets provide more uniform heat transfer distributions as compared with the normal jets. For the normal jet impingement and the tangential jet impingement, the RKE model provides better prediction than the SST model. The results can be useful for selecting a jet impinging position in order to provide the proper cooling distribution inside a turbine blade leading edge region.

An experimental investigation of cooling characteristics at a vane end-wall with a locally enhanced hole-layout

This paper presents an experimental investigation on the effect of film hole-layout at a turbine vane end-wall on the cooling performances, including the distributions of overall cooling effectiveness, uniformity of temperature field and highest temperature drop, using a hot wind tunnel and an Infrared Thermal Imaging System (ITIS). Two specimens were manufactured and used. One is named as “uniform hole-layout (UHL)”, where ten rows of film holes were uniformly arranged along the axial direction. However, UHL results in a non-uniform temperature distribution over the entire end-wall surface, and the low film coverage and bad reliability in the regions near leading edge (LE) and pressure side (PS) of the vane thereby. To improve the cooling performance in these regions, a new layout named as “locally enhanced hole-layout (LEHL)” was proposed. The difference from the previous experiments conducted under adiabatic assumption is that, the conjugate heat transfer is considered, and the conjugate-temperature distribution is used as the basis of design of LEHL. The comparisons of the cooling performances of two specimens at various mass flow ratios (MFRs) of coolant-to-mainstream revealed that: (1) from MFR = 0.5% to 1.5%, using LEHL, the area-averaged overall cooling effectiveness over the entire end-wall, in LE region and in the corner of PS can be increased by at least 32.3%, 34.2% and 26.2%, respectively; and (2) the temperature differences from PS to suction side (SS) and the temperature at the entire end-wall can be significantly reduced by LEHL. Therefore, the design strategy of new hole-layout using the conjugate-temperature is conducive to obtain the simultaneous improvement of film coverage and working life.

Multi-step optimizations of leading edge and downstream film cooling configurations on a high pressure turbine vane

Multi-step optimizations of leading edge and downstream film cooling configurations on a high pressure turbine vane are conducted using full three-dimensional optimization method. Firstly, a full three-dimensional optimization platform for film-cooled turbines, which consists of the parametric modeling, automatic mesh generation, the CFD numerical calculation and the optimization strategy is established. On the basis of it, the optimization platform is used to optimize the configurations of multi-rows film cooling holes for C3X gas turbine vane cooled with leading edge showerhead film cooling, and the layout of rows of film cooling holes on pressure side and suction side is optimized for the further step optimization. The sensitivity analysis for film cooling performance is used to decide the final design variables, and the adiabatic film cooling effectiveness is adopted to evaluate the overall performance of film-cooled turbine vane. The results show that the average film cooling effectiveness of the optimized leading edge region, pressure surface and suction surface are increased by 8.4%, 1.1% and 5.5% compared with the original case, while the average temperature of the leading edge region is decreased by 19 K.

Effect of Thermal Conductivity on Nozzle Guide Vane Internal Surface Temperature Distribution

Abstract Conjugate heat transfer analysis is carried out in a cascade domain for a nozzle guide vane. The nozzle guide vane is internally cooled by jet impingement cooling, and the external surface is cooled by film cooling. A computational study was carried out with three different materials, having conductivity values of 0.0048, 0.2 and 1.1 W/m.K. Distribution of local surface temperature along the leading edge, pressure and suction surface is reported. The leading edge region showed the maximum increase in internal surface temperature as the conductivity increased among the different regions of the vane internal surface. However, the pressure and suction surfaces showed relatively less increase in the surface temperature distribution. In order to validate the computational result, the obtained temperature data were compared with experimentally obtained surface temperature data. The flow phenomena like jet lift-off and self-induced cross-flow affect the local temperature distribution differently in the three materials. For a constant mainstream and coolant flow, the surface temperature gradient is higher for the lower conductivity material, and the gradient decreases as conductivity increases. Hence, a material with higher conductivity is desired in a combined impingement and film cooled nozzle guide vane, to increase the durability of the vane.

Reduced innate immunity of Cuban Treefrogs at leading edge of range expansion.

During geographic range expansion, populations of non-indigenous species at the invasion front may benefit from directing resources away from immune defense. To test this hypothesis, we investigated the strength of two innate immune components in populations of invasive Cuban Treefrogs (Osteopilus septentrionalis) in a long-colonized area (core region) and at the invasion front (leading-edge region). First, we compared the region-specific metabolic response of frogs injected with an endotoxin that induces systemic inflammation (lipopolysaccharide, LPS) to sham-injected control frogs pooled from both regions. Males and females were analyzed independently because we detected a sex-related difference in mass-independent metabolism of control frogs, with males exhibiting a significantly higher metabolic rate (F1, 21 =29.02, P<0.001) than females. We observed a significantly higher metabolic rate in LPS-injected core frogs compared with control frogs for both males (P = 0.041) and females (P = 0.007). Conversely, in leading-edge populations, there was no significant difference in the metabolic rate of LPS-injected and control frogs (males, P=0.195; females, P=0.132). Second, we directly compared bacterial killing ability of frog blood plasma between regions. Bactericidal ability of plasma was significantly greater in frogs from the core region in comparison with those at the leading edge (F1, 26 =28.67, P<0.001). For both immune components that we examined, populations from the core exhibited stronger immune responses. Our findings support hypotheses predicting an inverse relationship between immunity and range expansion.

Film Cooling Performances at Different Turbulence Intensities Using Conjugate Heat Transfer Analysis

This study presents a numerical investigation of cooling performances of a modified vane of the film-cooled vane reported by Timko (NASA CR-168289) at different mainstream turbulence intensities (Tus). A 3D conjugate heat transfer (CHT) analysis with SST k-ω turbulence model in FLUENT V.15 is used. Three different mechanisms in CHT analysis, i.e. fluid flow, heat convection between solid surfaces and flowing fluid in an external mainstream and internal cooling passages, and heat conduction within the vane structure, are simultaneously considered. Numerical results are conducted in terms of overall cooling effectiveness at Tu=3.3, 10, and 20%. Comparison between overall cooling effectiveness and film effectiveness under adiabatic assumption is discussed at the three Tus, also. The findings of this research indicate the following phenomena: 1) overall cooling effectiveness decreases with Tu, and this effect on the pressure side (PS) is stronger than that on the suction side (SS) in general. 2) By comparison with adiabatic film effectiveness, the level of overall cooling effectiveness in most regions is higher and more uniform than that of adiabatic film effectiveness for all three Tus. 3) In the leading edge (LE), when Tu increases, near the exits of film holes overall cooling effectiveness deteriorates, but adiabatic film effectiveness improves. Furthermore, a large area with relatively low overall cooling effectiveness is able to move with Tu in the LE region.