Endwall Surface Research Articles

Effects of Burst Ratio and Frequency on the Passage Vortex Reduction of a Linear Turbine Cascade Using a Dielectric Barrier Discharge Plasma Actuator

A dielectric barrier discharge plasma actuator was employed to reduce the passage vortex generated in a turbine cascade. This study focused on the burst mode drive of a plasma actuator and examined the relationship between flow field changes and the burst ratio and frequency. The non-dimensionalized burst frequency was fixed at F+ = 1.26, and the burst ratio was varied from 0.01 (1% operation) to 1 (100% operation, continuous mode). Generally, an increase in the burst ratio weakens the passage vortex, and the center of the passage vortex moves more toward the upper endwall surface and blade suction surface side. However, the velocity distribution, secondary flow streamlines, turbulence intensity distribution, and vorticity distribution did not change proportionally with changes in the burst ratio. Furthermore, the burst ratio was fixed at BR = 0.5, and the non-dimensional burst frequency varied from F+ = 0.013 to 62.9. Low burst frequencies led to a decrease in the peak velocity of the passage vortex, vorticity at the passage vortex center, and negative peak vorticity; however, an increase in the distance of the center of the passage vortex from the upper endwall surface and the turbulence intensity at the vortex center was observed. In contrast, high burst frequencies resulted in a decrease in the position of the vortex center and the turbulence intensity of the passage vortex center, while the peak velocity of the passage vortex, vorticity at the vortex center, and negative peak vorticity increased. The non-dimensionalized burst frequency around F+ = 1 is appropriate because both effects are balanced.

The Investigation of a New End Wall Contouring Method for Axial Compressors

To further control corner separation in high-load axial compressors, this study proposes a new end wall contouring method. It defines multiple standard “surface units” with particular flow control effects and then applies a linear combination, finally forming the geometry of the end wall surface. Based on design experiences, three different end wall contouring cases are generated and calculated on a high-load compressor cascade in the first step. The results show that the new method achieves a clear and intuitive influence on the end wall geometry, with a proper number of design variables, and can effectively combine variables with the development of secondary flow. In the second step, the new method was applied to an axial compressor, with an improvement in the design variables. Although the end wall contouring only improved the efficiency of the compressor stage on the right part of its operating map, the experimental results of the flow field show that the corner separation and end wall loss are suppressed at multiple inflow conditions. The results thus verified the practical effect of the newly developed end wall contouring method.

Numerical Investigation into a New Method of Non-Axisymmetric End Wall Contouring for Axial Compressors

To deal with corner separation in high-load axial compressors, this paper proposes a new end wall contouring method aimed at controlling the end wall secondary flow in more than one local area, generating a geometry with fewer control variables that is applicable for multiple working conditions. The new method defines more than one surface unit function, with different effects on end wall secondary flow. Then, the geometry of these surface unit functions will be superposed to generate the end wall contouring, to combine their flow control effects. After applying the new method to a bi-objective optimization design process, with 15 design variables aimed at minimizing the loss of cascade at 0° and 4° incidence, the optimal design reduces the total pressure loss of the high-load cascade by 5% under the former incidence and by 3% under the latter. The most effective design rule is constructing an end wall surface with the rising suction side and sinking pressure side in the blade channel, while locally raising the SS corner with a gentle upstream slope. According to the analysis, the design variables of the new method show an intuitive influence on the variation of end wall geometry and the movement of secondary flow. The corner separation has been effectively suppressed, with fewer control variables than before. It, thus, indicates the advantage of the newly developed end wall contouring method compared with previous studies.

Effects of Near-Wall Streamwise Vortex on the Film Cooling Performance of a Fan-Shaped Hole

Abstract Fan-shaped hole is a typical shaped hole frequently used in gas turbine film cooling, which may operate in a complex vortical flow environment of turbine blade rows. It is critical to understand the destructive mechanism of the film cooling by near-wall vortex on the endwall surface. A vortex generator (VG) was installed in front of a discrete fan-shaped hole to produce a vortical environment on a flat plate. Heat transfer effects were experimentally studied for mainstream velocities of 20 m/s by measuring the adiabatic film cooling effectiveness at blowing ratios M = 0.5 to 2.5. Fluid dynamics effects of vortical upwash and downwash movements were numerically analyzed. Results show that near-wall streamwise vortex destroys the film flow and intensifies the mixing of hot gas and coolant flow, which leads to the deterioration of film cooling performance. Film coverage area varies remarkably with respect to the positions of streamwise vortex. An efficient blowing ratio has a strong ability to mitigate the influence of near-wall vortex. At M = 2.5, streamwise vortex can restrain liftoff tendency of coolant jet and retard the decrease of film cooling effects.

Experimental study of cooling air effect on overall cooling of laminated configuration at a turbine vane end-wall

Turbine vane end-wall cooling management is complicated by the sophisticated cascade flow, generating the non-uniform distribution of metal temperature at end-wall surface, a negative factor of component durability. The double-wall laminated cooling has a huge potential to achieve the two-side thermal protection for end-wall, as well as the minimum cooling air amount. However, the design guidance of laminated cooling end-wall is still scarce, since that the reported flat-plate studies ignored the influence of end-wall flow. In present work, the cooling air effect on the metal overall effectiveness was extracted from the infrared thermography measurements under different cooling air flowrates, indicating that there is an utmost of cooling air consumption for the reduction of metal temperature. The two-dimensional contour of effectiveness induced by cooling air was selected as the new design principle. A case study indicated that the modified laminated cooling layout can acquire the uniform distribution of metal temperature over the entire end-wall surface. A comparison of metal overall effectiveness of original and modified end-walls revealed that in leading edge of pressure side, the novel film cooling design is more proper; and the dense layout of impingement-hole with large diameters is effective in the corner of pressure side.

Study on film cooling for aero-engine turbine endwalls with various small-scale surface structures

In this study, a novel endwall modification by adding small-scale ribs onto a turbine endwall surface is proposed to reduce thermal loads on the endwall. The small-scale ribs are positioned in a streamline-curved and an axially-straight array patterns and, square and triangular shapes are imposed to the cross section of the ribs. The feasibility of improving purge air cooling by the ribs is investigated by using well-validated computational fluid dynamics (CFD) methods. Comprehensive comparisons of cooling performance among the ribbed endwall cases are made by examining cooling effectiveness, heat transfer, and aerodynamic losses. Detailed flow topologies are presented to reveal flow mechanisms involved in the actions of the ribs within the turbine passage flow. Particularly, secondary-order effects of the ribs on the turbine airfoil surfaces are also documented, which is expected to provide three-dimensional boundary conditions for the turbine aerodynamic and thermal analyses. Comparisons between the ribbed endwall cases and a smooth endwall case show that the endwall modification by adding small-scale ribs is effective in reducing thermal loads on the endwall via enhancing cooling effectiveness and, in some cases, reducing heat transfer levels. The thermal load reduction can be further improved by increasing purge air flow ratio. However, as additional vorticities are introduced by the ribs, slight aerodynamic penalty is generated. Among the four rib cases, the straight triangular-shaped rib case has the lowest overall cooling performance for the endwall. Additionally, secondary-order cooling effects are visible on the airfoil suction surface for all rib cases but only the straight rib cases can generate additional purge air coverage on the airfoil pressure surface, without significantly changing pressures in that region. This work offers a new option of thermal management techniques for the turbine endwall and the results provide guidelines regarding endwall modifications of such sort.

Cooling holes with different structures effects on coolant cross flow and heat transfer at the outlet of the combustion chamber

Abstract: The major effects of cylindrical and row trenched cooling holes with angles of alpha=30, beta=0, alpha=40, beta=0 and alpha=50, beta=0 at BR=3.18 on the effectiveness of film cooling near the combustor end wall surface is an important subject to study in detail. In the current study, researchers used a FLUENT package 16/11 to simulate a 3-D model of a Pratt and Whitney gas turbine engine. In this research, RNG turbulence model K-ε model was used to analyze the flow behavior on the passage ways of internal cooling. In the combustor simulator, the dilution jets and cooling flow staggered in the streamwise direction and aligned in the spanwise direction as well. In comparison with the baseline case of cooling holes, the application of trenched hole near the end wall surface increased the effectiveness of film cooling up to 100% for different trench cases.

확장형 막냉각 홀이 적용된 터빈 블레이드 끝벽 막냉각 효율 측정

In this study, the adiabatic film cooling efficiency measurement using pressure-sensitive paint under conditions simulating the actual turbine blade endwall environment was conducted to verify that the optimal fan-shaped film cooling hole developed through the plate film cooling experiment was effective even in an actual turbine environment. Two end wall specimens with the baseline fan-shaped film cooling hole and the optimal fan-shaped film cooling hole were fabricated and compared the adiabatic film cooling efficiency at the endwall surface of the turbine blade. The optimal fan-shaped film cooling hole was able to effectively cool a wider area with the effect of widening the expansion angles in both the forward and lateral directions compared to the baseline fan-shaped film cooling hole. As a result, the turbine blade endwall showed different film cooling efficiency characteristics than that of the flat plate due to the pressure gradient generated by the turbine blades. However, as in the previous study results, the optimal expansion type film cooling hole showed better film cooling efficiency under all conditions, and the difference became larger as the film cooling flow rate increased.

Effective strategy of non-axisymmetric endwall contouring in a linear compressor cascade

The corner separation and the related secondary flow have great impact on the compressor performance, and non-axisymmetric endwall contouring is proved effective in improving compressor efficiency. The aim of the study is to improve the compressor performance by two local endwall contouring strategies at the design and off-design conditions. The endwall is parameterized and the Bezier curve is used to loft the endwall surface. The design of the contoured endwall is based on a multi-point optimization method to minimize the aerodynamic pressure loss. In order to identify the influence of the contoured endwall, a detailed flow analysis is conducted on four effective contoured endwall designs. The selected endwall geometries exhibit great control ability on the corner separation and significantly reduce the pressure loss at the two operating conditions. The directional concave near the leading edge can induce strong streamwise pressure gradient and accelerate the endwall flow, greatly reducing the cross-passage pressure gradient. The convex structures near the concave edge and at the outlet can block the cross-flow and prevent the interaction between the cross-flow and the suction corner flow. The benefit of the contoured endwall is mainly due to the re-distributed endwall static pressure and blocking of the cross-flow movement. In terms of the control effect, the shape of the concave also matters and better control effect is observed on the deep and wide concave. The flow will be guided by the concave, and the best suppression on corner separation is observed on the concave which follows the suction side. The results also indicate that the relief of the hub corner separation slightly increases the shroud pressure loss.

On the unsteady behaviours of the adiabatic endwall film cooling effectiveness

The film cooling technique is introduced in modern gas turbines to protect the blade from the high temperature of the incoming hot gases by forming a thin coolant blank over the blade surface. However, it is known as a jet in crossflow (JICF), where coolant and mainstream interact intensively and generate complex vortices leading to highly unsteady coolant coverage over the blades surface. In this study, a fast-response pressure-sensitive paint technique (fast-PSP) was used to measure the coolant unsteadiness with a high-resolution camera. The measurements were performed in a novel single-passage transonic wind tunnel to uncover the unsteady effectiveness of the endwall surface. Such effectiveness was dramatically influenced by the blowing ratios (M), showing attached flow at a low blowing ratio and lift-off at a high blowing ratio. The effectiveness was asymmetrically distributed due to the pressure gradients, jet compounding angle, and associated complex flows. The unsteady effectiveness was highly influenced by the energetic vortical structures, which interacted with the mainstream flow immediately behind the holes. It was featured by secondary structures (horseshoe, passage, and counter vortices) beside the JICF structures. Meanwhile, the unsteadiness was originated from the middle of the passage behind the holes. It is suggested to pay close attention to the locations of the holes for further optimization. This study could help the designers to understand the characteristics of unsteady effectiveness, promoting advanced cooling strategies for enhanced protection of future gas turbines.

Investigations of improved cooling effectiveness for ramp film cooling with compound angle film cooling jets

PurposeThis paper aims to investigate the film cooling effectiveness (FCE) and mixing flow characteristics of the flat surface ramp model integrated with a compound angled film cooling jet.Design/methodology/approachThree-dimensional numerical simulation is performed on a flat surface ramp model with Reynolds Averaged Navier-Stokes approach using a finite volume solver. The tested model has a fixed ramp angle of 24° and a ramp width of two times the diameter of the film cooling hole. The coolant air is injected at 30° along the freestream direction. Three different film hole compound angles oriented to freestream direction at 0°, 90° and 180° were investigated for their performance on-ramp film cooling. The tested blowing ratios (BRs) are in the range of 0.9–2.0.FindingsThe film hole oriented at a compound angle of 180° has improved the area-averaged FCE on the ramp test surface by 86.74% at a mid-BR of 1.4% and 318.75% at higher BRs of 2.0. The 180° film hole compound angle has also produced higher local and spanwise averaged FCE on the ramp test surface.Originality/valueAccording to the authors’ knowledge, this study is the first of its kind to investigate the ramp film cooling with a compound angle film cooling hole. The improved ramp model with a 180° film hole compound angle can be effectively applied for the end-wall surfaces of gas turbine film cooling.

The Role of Secondary Flows and Separation in Convective Heat Transfer in a Rotating Radial Vane Brake Disk

Abstract This study presents, for the first time, distributions of local internal temperature and convective heat transfer in a rotating radial vane brake disk and explains mechanisms in conjunction with secondary flows and flow separation within its ventilated coolant passages. In particular, variations of radial, circumferential (vane-to-vane), and axial (inboard-to-outboard) heat transfer on internal end-wall surfaces, and their alteration due to varying number of radial vanes and rotating speed are experimentally detailed. It has been demonstrated that conventional ventilated radial brake disks where the air inflow is drawn from the inboard face are likely to suffer substantial axial variations of temperature and heat transfer between the inboard and outboard disks, which possibly exacerbates thermal distortion (i.e., coning). Further, for a typical number of vanes (i.e., 36 vanes) used on automobiles, internal thermal distributions are highly nonuniform. However, the thermal end-wall uniformity improves considerably as the number of vanes is increased to say 72 vanes. Specifically, as the number of vanes is increased, secondary flow mixing enhances overall convective heat transfer and improves thermal uniformity. In contrast, separation causes large end-wall thermal nonuniformities in radial and circumferential distributions between the pressure side and the suction side of radial vanes. This effect nonetheless also decreases as the number of vanes is increased.

Implementation of Rectangular Vortex Generator Pairs to Improve Film Cooling Effectiveness on Transonic Rotor Blade Endwall

Abstract The rectangular vortex generator pairs (RVGPs) are arranged upstream of the film cooling holes to achieve better coolant coverage on the endwall near the pressure-side corner area. The endwall film cooling effectiveness distributions under transonic flow conditions are numerically calculated for the single RVGP and double rows of RVGPs cases. At first, the effects of three geometrical parameters (i.e., the distance between RVGP and cooling hole, height of RVGP and attack angle of RVGP) on endwall film cooling effectiveness are studied with a single hole and RVGP at different mainstream inlet Reynolds numbers and blowing ratios. Then, the double rows of RVGPs are applied to further enhance the overall film cooling effectiveness on the blade endwall. The results show that the implementation of RVGPs significantly enhances the film cooling effect on transonic blade endwall at a pressure-side corner area. With the increase of RVGP height, the lateral coolant coverage on the endwall corner area is improved. However, by decreasing the distance between the vortex generator pair and cooling hole, the film cooling effectiveness downstream of the cooling holes is increased. The attack angle of RVGP mainly affects the shape of coolant spreading on endwall surface. The RVGP with optimum dimensions and arrangement is able to suppress the coolant from lifting off the endwall and increase the coolant diffusion near the endwall. Compared with no vortex generator case, the area-averaged film cooling effectiveness on endwall with double rows of RVGPs is improved by 13.16%.

Nozzle Passage Endwall Effectiveness Values With Various Combustor Coolant Flowrates—Part 2: Endwall and Vicinity Surface Effectiveness Measurements

AbstractEffective coolant schemes are required for providing cooling to the first-stage stator vanes of gas turbines. To correctly predict coolant performance on the endwall and vane surfaces, these coolant schemes should also consider the effects of coolant streams introduced upstream in the combustor section of a gas turbine engine. This two-part paper presents measurements taken on a first-stage nozzle guide vane cascade that includes combustor coolant injection. The first part of this paper explains how coolant transport and coolant-mainstream interaction in the vane passage is affected by changing the combustor coolant and endwall film coolant flowrates. This paper explains how those flows affect the coolant effectiveness on the endwall and vane surfaces. Part one showed that a significant amount of coolant injected upstream of the endwall is present along the pressure surface of the vanes as well as over the endwall. Part two shows effectiveness measurement results taken in this study on the endwall and pressure and suction surfaces of the vanes. Sustained endwall coolant effectiveness is observed along the whole passage for all cases. It is uniform in the pitch-wise direction. Combustor coolant flow significantly affects cooling performance even near the trailing edge. The modified flowfield results in the pressure surface being cooled more effectively than the suction surface. While the effectiveness distribution on the pressure surface varies with combustor and film coolant flowrates, the distribution along the suction surface remains largely unchanged.

Nozzle Passage Endwall Effectiveness Values With Various Combustor Coolant Flowrates—Part 1: Flowfield Velocity and Coolant Concentration Measurements

Abstract The stators of the first stage of a gas turbine are exposed to severe temperatures. The coolant streams introduced to prevent the stators from thermal damage further complicate the highly three-dimensional vane passage flow. Recent results have shown that the coolant streams injected for cooling the combustor also influence the flow physics and the cooling effectiveness in the first-stage stator vanes passage. However, the effects of changing the mass flowrate of these combustor coolant streams on the passage flowfield have not been studied. As understanding the coolant transport is necessary for analyzing changes in cooling effectiveness in the vane passage, detailed aerodynamic and thermal measurements along the whole vane passage are required. This two-part paper presents such measurements taken for a variety of combustor coolant and endwall film coolant flowrates. The experiments were conducted in a low-Mach number facility with engine-representative Reynolds numbers and large-scale high-level turbulence. The objective of the first part is to describe the flow that influences endwall and vane surface cooling effectiveness distributions, which are presented in the second part. The measurements show changes in the passage flowfield due to changes in both combustor coolant and endwall film coolant flowrates. Overall, the flow physics remains largely unaffected by changes in coolant flowrates except in the endwall-vane surfaces region where the combustor coolant flowrate dominates changes in coolant transport. This is shown to have a high impact on endwall and vane surface cooling.

Heat transfer enhancement and flow characteristics in a rectangular channel having inclined pin arrays mounted on the endwall surface

This study describes heat transfer and flow characteristics for various angles of inclined-pin staggered arrays in a rectangular channel. The pin array consists of 10 staggered rows of cylindrical pins. Influence of the pin inclined angle is studied at θ = 30o, 45o, 60o, 90o, 120o, 135o, and 150o. A thermochromic liquid crystal sheet is used for temperature measurement. Flow field and heat transfer characteristics are predicted by numerical simulation. From the study, certain pin inclined angles have shown a favorable effect on the heat transfer enhancement with an array of pins. Only pin arrays with inclined angles at θ = 120o and 135o could enhance heat transfer on the surface when compared with conventional right-angle pin. Jet-like flows behind the pin move the counter-rotating vortices with high velocity (CRH) that induce airflow attachment to the heat transfer surface down to the lower surface. So, CRH has a significant effect on heat transfer promotion. Pin inclination angles at θ = 120o, 135o and 150o could greatly reduce the pressure drop and enhance the thermal performance factor. The greater number of rows in the test section has been proven crucial for some inclination angles in promoting heat transfer performance.

Aerodynamic performance of profiled endwalls with upstream slot purge flow in a linear turbine cascade having pressure side separation

In aeroengines, purge flow directly fed from the compressor (which bypasses the combustor) is introduced through the disk space between blade rows to prevent the hot ingress. Higher quantity of purge gas fed through the wheel space can provide additional thermal protection to the passage endwall and blade surfaces. However, the interaction of purge flow with the mainstream flow leads to higher secondary losses. Secondary losses inside a turbine blade passage can be reduced effectively by endwall contouring. This paper presents computational investigation on the influence of non-axisymmetric endwall contouring over endwall secondary flow modification in the presence of purge flow with the pressure side bubble (PSB). The experimental analysis was conducted for the base case without purge and base case with purge (BCP) configurations having flat endwalls. The total pressure loss coefficient and exit yaw angle deviation were measured with the help of a five-hole pressure probe. Static pressure distribution over the blade midspan was obtained by 16 channel Scanivalve. Aerodynamic performances of three different profiled endwalls are numerically analyzed and are compared against the BCP configuration. The effects of different contoured endwall geometries on endwall static pressure distribution and secondary kinetic energy were also discussed. Analysis shows that in the first contoured endwall configuration (EC1), the formation of stagnation zones at a contour valley close to the suction surface causes the exit total pressure loss coefficient to increase. The shifting of the contour valley near to the pressure surface (EC2 configuration) has resulted in local acceleration of the diverted pressure side leg of the horseshoe vortex over the hump toward the end of the passage. In the third configuration (EC3 configuration), reduced valley depth and optimum hump height have effectively redistributed the endwall pitchwise pressure gradient. The increased static pressure coefficient at the endwall near to the pressure surface has eliminated the PSB formation. In addition, computational results of unsteady Reynolds averaged Navier–Stokes simulations are obtained for analyzing transient behavior of PSB, with more emphasis on its migration on the pressure surface and transport across the blade passage. The additional work done by the mainstream fluid to transport the low momentum PSB fluid has caused higher aerodynamic penalty at the blade exit region. In this viewpoint, the implementation of contoured endwalls has shown beneficial effects by eliminating the PSB and related secondary vortices. At 27% of axial chord downstream of the blade trailing edge, a 4.1% reduction in the total pressure loss coefficient was achieved with endwall contouring.

가스터빈 연소기 출구 열·유동장 분포가 터빈에 미치는 효과의 수치해석 연구

Effects of temperature and flow fields at a combustor outlet on a high pressure gas turbine stage are investigated by using numerical methods. Numerical simulations of the annular type combustor are carried out by using large eddy simulation (LES) and unsteady Reynolds-averaged Navier-Stokes equations (URANS) techniques. The temperature distribution is found to be significantly affected by vortical structures at the primary zone and dilution jets. Numerical simulations of high pressure turbine stage are also conducted by using the Reynolds-averaged Navier-Stokes equations (RANS). Temperature distributions and aerodynamic loads at blade and end wall surfaces are drastically altered by changes in temperature, momentum, turbulent kinetic energy and turbulence dissipation rate at the exit of the combustor which are imposed as turbine inlet boundary conditions.

Endwall film cooling holes design upstream of the leading edge of a turbine vane

Endwall film cooling upstream of the leading edge (LE) of a vane presents a relatively complex flow phenomenon due to the horseshoe vortices (HVs) generated at the LE region. Upstream of the LE region, the coolant flow has difficulties to eject out. This research work focuses on controlling the jet holes coolant coverage upstream of the LE region. Compound angle holes in staggered arrangement are introduced and applied in the LE region to increase the coolant coverage. Six different arrangements of cooling holes are designed, with variations from one row or two rows arrangements, parallel or staggered arrangements, normal cylindrical holes or compound angle holes. Besides cooling holes with different shapes and arrangements, effect of the blowing ratio (BR) and turbulence intensity (TI) are also considered. The BR ranges from 1 to 3 and TI ranges from 1.3% to 15%. The calculated results show that the film cooling holes upstream of the LE of a vane have significant cooling effects on both the vane surfaces and the endwall. At small BRs, the film cooling effectiveness (η) on the endwall is considerable. When the BR is increased, the η on the vane surfaces is increased more quickly and becomes dominant. The coolant coverage in the vane-endwall junction region are not affected by the mainstream turbulence intensity and almost keep the same with the varied turbulence intensities. A single row of cylindrical holes (Case 1) and two rows of compound angle holes with staggered arrangement (Case 5) have relatively high overall averaged cooling effectiveness compared with other cases at different BRs. In addition, the high averaged cooling effectiveness on the endwall and vane surfaces by Case 5 is not affected by a change of the BRs.

Full coverage film cooling on a non-axisymmetric contoured endwall and the baseline endwall

Full coverage film cooling arrangement is widely needed in nozzle guide vanes (NGVs) to keep the turbine surfaces from melting due to the unprecedented high turbine inlet temperature. Meanwhile, to reduce the passage aerodynamic losses, non-axisymmetric endwall contouring are widely applied. The employment of endwall contouring will impact the full coverage film cooling performance, thus changing the turbine durability. Thus, it is important to deeper know the effect of non-axisymmetric endwall contouring on endwall full coverage film cooling performance. In this study, Pressure Sensitivity Paint (PSP) technique is used to measure the full coverage film cooling effectiveness (η) distributions on a non-axisymmetric contoured endwall (CE) and the corresponding baseline endwall (BE) with several coolant mass flow ratios (MFRs). Meanwhile, cases with the corresponding experimental conditions by computational fluid dynamics (CFD) are done to calculate the blowing ratio (BR) and MFR distributions of the film cooling holes. The η distributions on the CE and the BE are compared, and the effect of MFR on the CE full coverage film cooling performance are documented. In addition, different turbulence models are evaluated on the ability to model the endwall film cooling problems by comparing the CFD results with the experimental results. With small MFR values, the film cooling performance are worse on the CE in the near-hole regions at the fore part of the passage but better on the CE in the far-hole regions and at the aft part of the passage than that on the BE. While with large MFRvalues, the CE perform better than the BE throughout the endwall surface.