Rows Of Film Holes Research Articles

Effects of Ribbed Crossflow Channel in a Turbine Blade on Film Cooling Performance of Diffusion Slot Holes with Various Cross Section Orientations

Abstract This paper investigates the influence of ribbed crossflow on the film cooling performance of a turbine rotor blade. A pressure-sensitive paint measurement technique was employed to measure the effectiveness of film cooling. The discharge coefficients were also measured to determine the flow resistance. A row of film holes was positioned at the pressure surface or suction surface with a spanwise hole spacing of 7.5D, which is half of the rib spacing. The experiments were carried out at a mainstream Reynolds number of 520,000, a turbulence intensity of 3.6%, and a density ratio of 0.97. The crossflow inlet velocity was 45% of the cascade inlet velocity. A fan-shaped hole with a 14 deg expansion angle (Fans-14), a horizontally oriented slot cross section diffusion hole with a 14 deg expansion angle (H1.7-14), and two vertically oriented slot cross section diffusion holes with 14 deg (V1.7-14) and 20 deg (V1.7-20) expansion angles were tested with/without crossflow. The results indicated that the slot cross-sectional orientation significantly changes the flow patterns inside the holes. H1.7-14 has stronger lateral expansion and better surface adhesion, while V1.7-14 and V1.7-20 yield more uniform lateral velocity distributions. Regardless of crossflow, H1.7-14 produces the highest film effectiveness and discharge coefficient on the pressure surface, while it changes to V1.7-20 on the suction surface. The ribbed crossflow increases the film effectiveness on both the pressure surface and suction surface, except for V1.7-20, as it is almost unaffected by the crossflow.

Investigation of the Discharge Coefficient for Laidback Fan-Shaped Holes on Turbine Blades Under Internal Crossflow Condition

Abstract The discharge coefficient of laidback fan-shaped holes on turbine blades under internal crossflow conditions is investigated using experimental and numerical simulation methods. The study is conducted on a linear cascade, with single rows of film holes set on both the pressure and suction surfaces of the test blade, coolant supplied by an internal crossflow channel. The internal crossflow Reynolds number ranged from 39,100 to 78,200, and the pressure ratio ranged from 1 to 1.5. This research considered the effects of internal crossflow Reynolds numbers and hole length-to-diameter ratio on the discharge coefficient of laidback fan-shaped holes. The findings are as follows: internal crossflow increases the inlet loss of film holes and flow separation within the holes, leading to a significant reduction in the discharge coefficient of film holes. Meanwhile, film holes on the pressure surface of the blade exhibit greater sensitivity to crossflow. Under different hole length-to-diameter ratio conditions, the length of the cylindrical section is the primary factor affecting the discharge coefficient, and the magnitude of the discharge coefficient depends on the extent of flow separation within the cylindrical section. The sensitivity of the discharge coefficient to variations in the cylindrical section length decreases once the cylindrical section length exceeds 2.8D for the pressure surface and 2.0D for the suction surface, where the inlet separation region reattaches to the wall. Under plenum conditions, there are significant differences in the discharge coefficients of film holes between the suction surface and pressure surface, especially at pressure ratios below 1.05.

Experimental and numerical investigations on film cooling performance of converging slot holes in plate and turbine conditions

Based on the heat-mass transfer analogy, the Pressure Sensitive Paint (PSP) technique was used to explore the film cooling characteristics of converging slot holes on the plate model and the actual turbine endwall model conditions. Meanwhile, the numerical calculation was also adopted to pursue the flow field details. The width of the converging slot hole exit is a key geometry affecting cooling performance. On the plate model condition, three widths of exit slot Lw (1.7d, 2.3d, 2.9d) were chosen to obtain the optimal exit width. The case with Lw = 2.3d obtained the highest level of adiabatic effectiveness from spanwise and streamwise views. This case operated as a two-dimensional slot and produced a more uniform cooling jet film relative to other cases. Nevertheless, the cooling performance decreased significantly as the converging slot hole width increased to Lw = 2.9d. Then the cooling performance of three patterns of film holes was further investigated in the turbine endwall. Laid-back fan-shaped holes and converging slot holes could improve the film cooling performance effectively compared to the cylindrical holes. Converging slot holes with Lw = 2.3d improved the uniformity of the cooling jet outflow through rows of upstream film holes, and also effectively avoided the phenomenon of coolant blowing away or high-temperature gas entering the film hole, resulting in high film cooling effectiveness performance. Compared to cylindrical holes, the area-averaged adiabatic effectiveness was increased by 216.5 % for the endwall with cylindrical holes with the momentum ratio I = 0.25. Overall, converging slot holes improved cooling performance among all studied film hole configurations.

Thermo-mechanical analysis and creep lifetime prediction of coated turbine vanes considering thermal barrier coating spallation characteristics

Quantitative analysis of the influence of thermal barrier coating spallation on the thermo-mechanical behaviors and creep lifetime of turbine vanes is crucial for their maintenance and reliability improvement. The temperature and stress distribution of vanes with preset coating spallation damage are investigated in this study, using the thermal-fluid-solid coupling method. Additionally, a further computational analysis is conducted to predict the hazardous regions and creep lifetime of the vanes. The results indicate that coating spallation leads to significant changes of the temperature and stress distribution at the spalled regions of vanes. The remaining coating on the unspalled regions continues to provide effective protection. Stress concentration primarily occurs at the upstream and downstream of the leading edge film holes, while high-stress regions are observed between adjacent rows of film holes, forming a serrated shape. The creep lifetime of the vanes decreases significantly at the region with coating spallation. When the same spallation area is considered, the coating spallation at the leading edge has a more serious influence on creep lifetime, which is more likely to cause the vanes failure. This study reveals the influence of coating spallation characteristics on the thermo-mechanical behaviors and creep lifetime of vanes, providing valuable insights for durability assessment of coated high-temperature components.

Effects of film hole spanwise angle and diameter on flow and heat transfer of impingement/effusion cooling with crossflow diverters

This paper studies the blade cooling under the co-effect of external film extraction and internal impingement cooling with crossflow diverters. The influence of spanwise angle (α) and diameter (Df) of film hole on flow and cooling performance are analyzed when the impingement jet Reynolds number ranges (Rei) from 15,000 to 45,000. From the results, if a row of film holes is set on the target surface of an impingement cooling with crossflow diverters, the overall cooling effectiveness will be greatly improved, especially the pressure side. When the α ranges from 35° to 115°, no matter pressure side or suction side, the film jet injection towards blade hub is more beneficial to cooling, and the smaller the α, the more beneficial. When the Df ranges from 0.25 mm to 1.00 mm, at suction side, as Df increases, overall cooling effectiveness gradually increases; at pressure side, when the Df ranges from 0.25 mm to 0.75 mm, overall cooling effectiveness increases, but if the Df continues to increase, it will decrease, which means the Df of 0.75 mm can obtain the most excellent cooling performance. This study provides a theoretical basis for the comprehensive application of film extraction, impingement cooling, and crossflow diverters in blade cooling design.

Film Cooling Performance on a Turbine Blade With Subregional Compound Angle

Abstract Under the influence of secondary flow, the film cooling deviates from the flow direction on the turbine blade, which directly results in undesirable uneven film coverage. On the pressure side, the film appears divergent, while on the suction side, it is bunched. To solve this problem, a kind of subregional compound angle is proposed, in which the angle in the spanwise direction is different in different regions depending on the strength and direction of the secondary flow. Four rows of film holes with five kinds of subregional compound angles are provided on the pressure side, while two rows of film holes with different subregional compound angles are provided on the suction side. The Reynolds Average Navier–Stokes (RANS) method of the SST k–ω turbulence model is chosen to solve the above blade arrangement. The results show that a significant improvement can be achieved by the introducing subregional compound injection of the film coolant compared to the case of simple injection. In compound injection, the injectant maintains sufficient momentum to prevent the coolant from being swept away by the secondary flow. This was found to be largely the case for most holes on the pressure side, and some holes on the suction side. However, for holes near the downstream section of the suction surface of the blade, where the passage vortex is strongest, no value is found for the compound angle that could redirect the coolant along the blade profile without radial deviation. In some cases, excessive values of the compound angle led to jet liftoff rather than spreading the film along the surface.

Film Cooling on Turbine Vane Endwalls With Different Inlet Cooling Configurations: Experimental and Computational Results

Abstract A comparative study of turbine endwall film cooling resulting from three different cooling configurations in front of the passage inlet was conducted. The three inlet cooling configurations considered in this study were a conventional continuous slot, double rows of discrete film holes, and an interrupted slot with a backward-facing step. Pressure sensitive paint (PSP) was sprayed over the endwall surfaces to map adiabatic cooling effectiveness contours and five-hole probe and thermocouple measurements were implemented to assess aero-thermal fields at the passage exit. The evolution of cooling effectiveness was compared across a full range span of coolant flow rates of 0.3−1.8% in a real engine. Additionally, complex flow structures inside the passage due to interactions of coolant with mainstream flows were visualized using computational fluid dynamics (CFD) simulations to support the observed coolant coverage patterns. A side-by-side comparison of CFD simulations against experiments was made as well to evaluate the reliability of conventional turbulence modeling methods in such a complex flow. The continuous slot was found to produce the highest cooling effectiveness values while the double-row holes offered the most uniform coolant coverage, leading to spreading out of the coolant jets downstream of the passage throat and even beyond the trailing edge of the passage for coolant injection rates higher than 1.0%. Supported by the CFD-predicted flow structures and measured aerodynamic losses, the injection from the double-row holes weakened the passage secondary flows and thus caused the lowest aerodynamic loss. On the contrary, the interrupted slot enhanced the secondary flows due to the backward step, resulting in the highest flow loss.

Numerical investigation on conjugate heat transfer characteristics of film and vortex composite cooling under rotating conditions

This paper numerically investigates the vortex and film composite cooling performance under rotating conditions. The cooling performance of the adiabatic and conjugate models is compared under the range of 0–4000 r/min. The conjugate model contains the fluid region: the cascade, the film holes, and the internal vortex cooling chamber, as well as the solid region: the blade material between the internal flow and the mainstream flow. The adiabatic model is established by removing the blade material part in the conjugate model. The dimensionless temperature θ inversely proportional to the temperature is adopted. Results show that the blade leading edge temperature doesn’t vary linearly with the rotating speed. The stagnation line of the mainstream flow on the blade leading edge moves from the pressure surface to the suction surface. The maximum θ appears at 1500 r/min when the stagnation line stays certainly on the row of film holes located on the pressure surface and is 7.86% higher than the minimum θ. The minimum θ appears at 2500 r/min when the stagnation line stays on the position between the rows of film holes. The distribution of θ is much uniform, and the value of θ is much higher in the conjugate cases than the adiabatic cases due to the heat conduction through the blade material. The highest aerodynamic parameter appears at 2000 r/min due to its relatively low blade leading edge temperature and low coolant consumption and is 41% higher than the aerodynamic parameter at 0 r/min.

EFFECT OF JET TO PLATE SPACING ON FILM COOLING PERFORMANCE IN A COMBINED IMPINGEMENT AND FILM COOLING ARRANGEMENT

The influence of jet to plate spacing on conjugate heat transfer from a flat plate with combined impingement and film cooling has been studied both numerically and experimentally. The flow configuration consists of a matrix of impingement holes and multiple staggered rows of cylindrical film holes. One side of the plate is impinged with air jets, while the other side is exposed to hot mainstream as well as the vented air from the film cooling holes. The effect of increasing jet to plate ratio (H/D = 0.4, 1.2, and 2.0) on film cooling flow structure and heat transfer is analyzed under conjugate boundary conditions. The flow and temperature patterns are predicted in the entire computational domain; namely, the plenum chamber, film holes, and mainstream-coolant interaction zone. The mainstream Reynolds number is varied from 49,050 to 130,800 while keeping the jet Reynolds number constant at 825 and the blowing ratio ranging from 0.6 to 1.6. The velocity, turbulence, and temperature distributions at the exit of the film hole together define the film coverage on the interaction surface. The thermochromic liquid crystal (TLC) technique is used to evaluate the surface temperature of the plate and validate the computed values of the surface effectiveness. Both measured and computed effectiveness distributions are found to be comparable. Increase in the H/D values beyond a certain limit causes decrease in film effectiveness on the interaction surface, especially downstream of the film hole. The influence of H/D is noticed for all the three blowing ratios investigated, albeit to varying degree.

Experimental and numerical study of swirl impingement cooling for turbine blade leading edge with internal ridged wall and film extraction holes

This paper presents a comparative experimental and numerical study of the internal heat transfer and pressure loss of swirl impingement cooling for turbine blade leading edge with internal ridges and film extraction holes. The leading edge cooling study includes a typical impingement cooling with centerline normal jets, a swirl impingement cooling with offset jets, and a swirl impingement cooling with offset jets and ridged wall. The normal jets are arranged at the impingement plate centerline, and the offset jets adopt a staggered arrangement alternatively on the lateral sides of the impingement plate. The ridged concave wall features two small semi-circular ridges symmetrically on opposite sides under the offset jets. Three rows of film holes are arranged symmetrically on the leading edge cavity, and another two rows of film holes are arranged on the opposite cavity wall behind. Transient thermochromic liquid crystal (TLC) experiments are conducted to obtain the detailed heat transfer distributions on the internal surface of the leading edge cavities with different cooling configurations. The results show that in the swirl impingement cooling, the swirling flow induced by the offset jets impinging onto the concave wall improves the streamwise heat transfer uniformity on the leading edge internal wall, which overcomes the limitation of large heat transfer non-uniformness between jet stagnation region and surrounding regions in typical normal jet impingement cooling. In the swirl impingement cooling with ridged wall, the near-wall jet interacts with the ridges on the concave wall and develops more widely along the spanwise directions, which greatly improves the spatial heat transfer uniformity. In addition, the ridges on the wall can also induce near-wall jets to separate from the wall and attach to the opposite concave wall, which produces low-velocity recirculation in the leading edge cavity and improves the capability and uniformity of coolant outflow through the film holes. Compared to the normal jet impingement cooling, the heat transfer of swirl impingement cooling with smooth and ridged wall can be increased respectively by 10.8%–14.0% and 10.9%–23.6% for the jet Reynold number from 20,000 to 50,000. In addition, the pressure loss is also reduced slightly by less than 5% in the swirl impingement cooling. As supplements to the experiments, numerical computations provide more insights into the turbulent flow structure in the leading edge cooling configurations.

Heat transfer and flow structure characteristics of film-cooled leading edge model with sweeping and normal jets

The heat transfer and flow structure characteristics of a film-cooled leading edge model with sweeping and normal jets were numerically investigated using the conjugate heat transfer method. The leading edge model is a simplified obtuse body model that has three rows of traditional cylindrical film holes, each with three holes inclined at 25° with respect to the surface and three angles oriented at 0° and ± 30° with respect to the axial direction. A normal jet has one row of cylindrical impingement nozzles, and a sweeping jet has only one fluidic oscillator. Three blowing ratios (M = 1.05, 2.07, and 4.11) and three temperature ratios (TR = 0.5, 0.625, and 0.75) are considered in the study. The shear stress transport k–ω turbulence model and an extremely fine structured grid were validated and applied to all simulations. The results indicate that the total pressure loss coefficient and heat transfer performance increase internally and externally, respectively, with the blowing ratio. However, they are insensitive to the temperature ratio in all cooling configurations. The sweeping jet and film (SJF) composite cooling exhibits 1.6162, 2.1195, and 2.1766 times the total pressure loss coefficient of the normal jet and film composite cooling when the blowing ratio increases from 1.05 to 4.11. However, the SJF offers the advantages of uniform Nusselt number distribution on the impinging surface and high overall cooling effectiveness on the external surface.

Numerical study of unsteady flow and cooling characteristics of turbine blade cutback trailing edges integrated with pin fins and film holes

Accurate prediction of the unsteady flow mixing process and its effects on cooling performance in the turbine blade trailing edge (TE) is essential to improving TE cooling designs. This paper used high accuracy transient Scale Adaptive Simulation (SAS) method to investigate the unsteady flow and cooling characteristics of two novel turbine blades cutback TEs under blowing ratios BRs=0.65 and 1.25. The two novel TEs have five rows of pin fins in the cooling slot and two rows of either cylindrical or fan-shaped film holes on the pressure side upstream of the cutback region. The results were compared against those of the baseline TE with only pin fins and the data obtained from Shear Stress Transport (SST) k-ω model.The results revealed that: 1) the coolant issued from upstream cylindrical or fan-shaped holes has significantly disturbed the Brown-Roshko vortices shedding off the lip of the top wall, but only largely reduced the magnitude of normal velocity fluctuation in the cutback region. The alternating vortices in the cutback region are stretched flat. The changes brought to the flow structures are more significant by the fan-shaped holes than the cylindrical holes; 2) Although increasing blowing ratio enhances flow unsteadiness in the cutback region of the three TEs, the differences in the magnitude of velocity fluctuations are much reduced among the three TEs; 3) The TE cutback with fan-shaped holes achieves a higher adiabatic wall cooling effectiveness ηaw on the cutback wall at BR=0.65, and also a lower reduction in ηaw at BR=1.25 than the TE cutback with cylindrical holes; 4) Transient SAS method demonstrates a better agreement of ηaw with the experimental data for the baseline TE on the cutback wall, and higher values of ηaw than the results of the SST k-ω model on the top wall.

Influences of Pressure ratio and Fluid Temperature on overall Cooling Performances of Ribbed Channel with Film Holes

The previous experiments of overall cooling performances were most conducted using simplified models and under the similar temperature ratio of mainstream to cooling air with real gas turbine operations, and ambient outlet pressure. To discuss the reliability of this type of experimental data, this paper exhibits two series of numerical simulations. Using a real E3 blade as model, which has two-pass rib-roughened channel with inclined film holes, numerical simulations are carried out at the same temperature ratio and pressure ratio, but different fluid temperatures including mainstream and cooling air, and different outlet pressure. The numerical results reveal two important conclusions: 1) At the same outlet pressure, the overall cooling effectiveness on PS is not sensitive to the fluid temperatures, but on SS in the region between two rows of film holes, a higher fluid temperature corresponds to a higher cooling effectiveness. 2) At the same pressure ratio of inlet to outlet, the overall cooling effectiveness on PS and SS is not sensitive to the outlet pressure and fluid temperature.

Effect of microchannel on combined impingement and film cooling of a concave surface

In this work, a micro-channel-based wall cooling approach is investigated. The cooling efficiency of curved surfaces is determined by employing a computational study for conjugate heat transfer. A semi-circular curved body depicting the turbine blade leading edge, with a microchannel placed in the solid wall of the body and three rows of film holes, is considered for the study. The film holes are positioned along a stagnation line and either side of the stagnation line at an angle of 25o. The steady-state solver with a realizable k-ε turbulence model is used to study the heat transfer results for film cooling in a 3D geometry. Simulations were carried out, and subsequently, parametric analysis was performed to observe the effect of varying blowing ratio and temperature ratio. The temperature distribution is observed to be more uniform due to the presence of the microchannel, resulting in a lesser thermal gradient on the curved surface. It is also noted that overall effectiveness increases with the blowing ratio. The maximum increase in overall effectiveness due to placing the microchannel is found to be ~16% for the blowing ratio of three.

Hole arrangement effect to film cooling performance on leading edge region of rotating blade

The effect of relative position of two rows of film holes on film cooling performance was experimentally and numerically investigated on each side (suction side and pressure side) of rotating blade leading edge. The film cooling performance was measured by pressure sensitive paint and phase-locked techniques. The numerical simulation was also employed to improve the understanding of flow filed. The first row of film holes is fixed at α1 = ±10° downstream of stagnation line and is followed by the second row of holes with staggered arrangement. The span-wise relative positions between two rows of holes, P2/P1, are ranged from 1/4 to 3/4. Meanwhile, the stream-wise location of second row of holes is varied from α2 = ±25° to ±40°. Five blowing ratios ranged from 0.5 to 2.0 are also employed. The results indicate that the effect of relative position of two rows on film cooling is different on the two sides of leading edge. On the suction side, the relative position between two rows of holes significantly affects the interaction between the adjacent jets from two rows, which determines the distribution and value of film cooling effectiveness. When it comes to the leading edge pressure side, closer distance between two adjacent holes in two rows produces better film cooling performance. Meanwhile, the effect of relative position between two rows of holes is mainly located in the near hole region.

Experimental Study on Film Cooling Effectiveness of Blade With Chevron-Shaped Holes Under Wake Influence

Abstract Gas turbines have been widely used. With the continuous improvement of the performance of gas turbines, the turbine inlet temperature has greatly exceeded the heat-resistance limit of the turbine blade material, so advanced cooling technology is required. The film cooling effectiveness distribution over the blade under the effect of wake was obtained by the pressure-sensitive paint (PSP) technique. The test blade has five rows of chevron film holes on the pressure side, three rows of cylindrical film holes on the leading edge, and three rows of chevron film holes on the suction side. The mainstream Reynolds number is 130,000 based on the blade chord length, and the mainstream turbulence intensity is 2.7%. The upstream wake was simulated by the spoken-wheel type wake generator. The film cooling effectiveness was measured at three wake Strouhal numbers (0, 0.12, and 0.36) and three mass flux ratios (MFR1, MFR2, and MFR3). The results show that the increase of mass flux ratio leads to a decrease of the film cooling effectiveness on the suction surface. In the wake condition, the effect of mass flux ratio is weakened. Wake leads to a marked decrease of the film cooling effectiveness over most blade surface except for the surface near leading edge on the pressure surface. In the high mass flux ratio condition, the effect of wake on the film cooling effectiveness is weakened on the suction surface and strengthened on the pressure surface.

Numerical investigation of blowing ratio, density ratio and axial position of film holes on the vane endwall film cooling effectiveness with upstream step

The aerothermal performance of interrupted slot and film holes was numerically investigated. Previous study indicates that the interrupted slot performs better compared to the conventional slot. In the meanwhile, the step formed along with the interrupted slot affects the film cooling characteristics. In this article, a row of film holes is arranged downstream of the step, and the mass flow rate for the interrupted slot is constant at 1%. Blowing ratio (BR) from 0.5 to 1.5 and density ratio from 1 to 2 were studied for the film holes. Endwall film cooling effectiveness distribution indicates that film cooling is easily affected by the secondary flow inside passage and the upstream step. Coolant traces are split into two parts due to the effects of step vortex and transverse flow. For different density ratios, increasing BR shows a different trend of film cooling effectiveness due to the variation of coolant momentum. The coolant jet is easily affected by the secondary flow when its momentum is low, but tends to liftoff when its momentum is too high. As a result, it is better to position the film holes far away from the upstream step. The total pressure loss coefficient distribution at the passage exit indicates that the coolant injection increases the total pressure loss. But density ratio has smaller effect on the loss variation. Besides, two axial positions of cooling holes were studied to improve the endwall cooling performance. Without the effect of step vortex, the film effectiveness of cooling holes is improved.

The cooling performance of multiple rows of film holes on the suction surface of a turbine blade under rotating conditions

In this paper, the experimental results on the cooling performance of multiple rows of holes on the suction surface of a turbine blade were indicated. The effects of hole arrangement, blowing ratio, and rotation on the film cooling performance were analyzed. Four different film hole arrangements were studied. Each blade was modeled with three rows of cylindrical holes with streamwise angle of 45°. The first and the second injection rows were fixed at the position of 12.4% and 17.8% of the streamwise length of the suction surface, respectively. The third injection row was located at streamwise location of either 23.2% or 34%. To investigate the rotation effect, experiments were performed at three rotational speeds of 400 rpm, 600 rpm, and 800 rpm corresponding to the rotational Reynolds number of 3.57 × 105, 5.36 × 105, and 7.15 × 105, respectively. Average blowing ratios of 0.5, 0.75, 1.0, 1.25, 1.5, and 2.0 were tested. Adiabatic film cooling effectiveness was obtained using pressure sensitive paint (PSP) technique. The results showed that inward deflection for the trajectories complicates the superposition principle. Staggered rows of holes can achieve more uniform film coverage and higher film cooling effectiveness. For the case with small array spacing between each adjacent rows, the film cooling area is too concentrated and can achieve better film cooling performance only at a small blowing ratio. The increasing rotational speed slightly weakens the tendency of inward film deflection and increases the film cooling effectiveness at a low blowing ratio.

Numerical investigation on flow and heat transfer characteristics of vortex cooling in an actual film-cooled leading edge

This paper numerically investigates the flow and heat transfer characteristics of vortex cooling in an actual film-cooled leading edge. Two internal vortex cooling cases called Case P and Case S are involved in this study. Case P means that the coolant injection nozzles are arranged near the pressure surface side of the blade, and Case S means that the coolant injection nozzles are arranged near the suction surface side. The mainstream region and the swirl chambers of the vortex cooling cases are connected by 6 rows of film holes. Results indicate that the practical mainstream flow generated by the actual blade affects the vortex flow structures and brings about an endwall secondary flow. For internal vortex cooling, the large-scale vortex flow structures in Case P skew off while flow structures in Case S are not distinctly influenced. Two type of flow regimes form near the film hole due to the suction effect, and bring about different local heat transfer augmentation intensity. For film cooling, the endwall secondary flow forces the coolant film move towards or away from the endwall thus decreases the film cooling efficiency. The film hole blowing ratio distribution is affected due to the different influence of mainstream flow on internal vortex flow of Case P and Case S, which results in different film cooling efficiency distribution. The film cooling efficiency of Case S on the pressure surface increases by up to 59.8% compared with Case P.

Experimental Study on Heat Transfer of Leading Edge Film-Cooling With Counter-Inclined Cylindrical and Laid-Back Holes

AbstractThe heat transfer coefficient of counterinclined film holes fed by different intake structures on the turbine vane leading edge (LE) model is experimentally investigated in this paper. A semicylinder model is adopted to model the vane leading edge, which is arranged with one single row of film holes per side, which are located from the stagnation at a 15-deg angle. The four leading edge models, which are the combinations of the hole-shapes (cylindrical hole and laid-back hole) and intake structures (plenum and impingement), are tested at four blowing ratios M. The contours of the heat transfer coefficient, which are characterized by the Frössling number Fr, since it includes the Reynold number effect, are acquired by the transient measurement technique based on double thermochromic liquid-crystals (LCs). The lateral-averaged Fr of the nonfilm-cooled model is provided by using the same experimental platform with an identical main-flow condition. It is then compared with the published data, which indicates the reliability of the present transient measurement techniques. The results illustrate that a core region with a higher heat transfer appears in the hole-exit downstream, and its distribution is slightly skewed to the inclination direction of the film holes. The shape of the high heat transfer region gradually inclines in the spanwise direction as M increases. The heat transfer in the region where the jet core flows through is relatively low, while the jet edge region is relatively high. The effect of impingement leads to the outflow of each hole becoming increasingly uniform, which can reduce the difference in the heat transfer between the region where the jet core flows through and the jet edge. The heat transfer strength may increase due to the intense turbulence caused by the introduction of the impingement. Compared with the cylindrical hole, the laid-back hole has a spanwise expansion feature, which makes the shape of the high heat transfer region wider in the spanwise direction and increases the heat transfer level. Additionally, the magnitude of the enhancement increases with an increasing M.