Fan-shaped Holes Research Articles

Large eddy simulation of in-hole blockage effects on cooling performance of fan-shaped holes

Film-cooling and thermal barrier coating approaches are commonly used in gas turbines to protect the turbine parts from the high-temperature flow of gases. However, one drawback of this method is that the coatings can partially obstruct the cooling hole exit on the turbine blade surfaces, resulting in reduced cooling performance. Therefore, this study aimed to investigate the effects of in-hole blockage thickness on cooling effectiveness and flow characteristics of fan-shaped holes using large eddy simulations (LES). A laidback fan-shaped hole located on a flat plate was the reference cooling hole. Numerical simulations were conducted for two different blocked cooling hole configurations with various blockage thicknesses (t/D = 0.5 and 0.9) at three different blowing ratios (0.5, 1.0, and 2.0). The cooling effectiveness computed by the LES method was validated for both unblocked and blocked cooling holes, and the results were compared to the experiments at various blowing ratios. It was revealed that an increase in blockage thickness had a significant impact on the area-averaged cooling effectiveness, particularly at high blowing ratios. The overall area-averaged cooling effectiveness for the t/D = 0.9 case was approximately 75 % lower than that of the unblocked cooling hole. Moreover, assessment of the instantaneous velocity field inside the hole over time revealed that larger blockage thicknesses resulted in greater velocity fluctuations and a more unsteady flow.

Large eddy simulations of the turbine vane pressure side film cooling flows of cylindrical and fan-shaped holes with a saw-tooth plasma actuator

Large eddy simulations were conducted to study the film cooling performance of turbine pressure side with cylindrical hole, fan-shaped hole and saw tooth plasma actuator (STPA). A detailed analysis of the time-averaged film cooling flows was made. The results showed that the dual control effects of the combined design of the fan-shaped hole with the STPA reduced the exit momentum and blowing off effect of the jet flow, remaining the jet flow close to the pressure side. The combined design also effectively weakened the entrainment effects of counter rotating vortex pair (CRVP) and enlarged cooling film coverage. Therefore, the spanwise averaged cooling efficiency was improved more than 30% in comparison to the cylindrical hole, but the fan-shaped hole caused a 23.9% increase in aerodynamic loss, and the STPA had few additional aerodynamic losses. Subsequently, the instantaneous film cooling flows were analyzed, the combined design suppressed the evolution processes of the vortex rings in the near-hole region, and the coherent structures in the far downstream region were decreased in size, affecting the mixing process of coolant and gases. The CRVPs were noticeably elongated along the pressure side and moved downstream periodically over time. The present study highlighted the superiority of the combined design to improve the turbine vane cooling performance.

Film Superposition Characteristics With Laidback Fan-Shaped Holes Under the Influence of Internal Crossflow

Abstract In the serpentine passage of actual rotor blades, the coolant at the inlet of the film hole has an internal crossflow. The internal crossflow significantly influences the film cooling effectiveness of multi-row film holes. This paper presents a numerical investigation of the superposition characteristics of multiple rows of laidback fan-shaped holes under the influence of internal crossflow. The numerical simulations utilize the RANS method considering the effects of blowing ratio, crossflow velocity ratio, arrangement patterns, and row-to-row spacing of the internal crossflow on the superposition characteristics of film holes. The results indicate that film cooling effectiveness exhibits a distinctly asymmetric distribution under crossflow conditions. At different blowing ratios, there exists an optimal crossflow velocity ratio, at which the laterally averaged film cooling effectiveness is maximized. The crossflow-to-jet velocity ratio (VRi) is utilized to comprehensively evaluate the impact of blowing ratios and crossflow velocity ratios, and the range of high area-averaged film cooling effectiveness is accurately captured. Regarding the study of arrangement patterns and row-to-row spacings, it has been found that under identical coolant mass flowrates, the staggered arrangement is better than the inline arrangement in both laterally averaged film cooling effectiveness and cooling uniformity. Meanwhile, reducing the row-to-row spacing leads to an improvement in laterally averaged film cooling effectiveness. Finally, an assessment of the Sellers model's applicability under crossflow conditions revealed that it demonstrates good applicability in cases of staggered arrangement with lower blowing ratios.

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.

A comprehensive comparison of film cooling characteristics of shaped holes on the leading edge of rotating blade

This study numerically investigates the cooling characteristics and flow features of various hole geometries on the rotating blade leading edge. A comprehensive comparison of cooling effectiveness, heat transfer coefficients, net heat flux reduction, and discharge coefficients is conducted among laidback fan-shaped, fan-shaped, laidback, cone holes, and cylindrical holes with two different diameters. Detailed flow and thermal field analysis reveals the impact of geometry on cooling performance. The blade rotational speed is 600 rpm, with a Reynolds number of 12.4 × 104, blowing ratios from 0.5 to 2.0, and fixed injection and spanwise angles at 90° and 30°. Shaped holes share a uniform hole exit-to-inlet area ratio, and cylindrical holes match their inlet or exit areas. The RANS equations were closed using the SST k-ω turbulence model coupled with the Gamma Theta transition model, and an unstructured grid of 13.0 million cells was employed to attain relatively accurate predictive results. Results show that the lateral expansion on the leading edge significantly enhances cooling performance, with fan-shaped holes outperforming others, especially at high blowing ratios. Additionally, cooling effectiveness and discharge coefficient on the leading edge correlate strongly with hole geometric parameters. Shaped holes with equal area ratios show similar discharge coefficients. Cooling effectiveness at high blowing ratios is proportional to the effective exit width, provided that the coolant flows smoothly without severe lift-off.

Effect and optimization of geometric parameters and arrangement on film cooling performance of fan-shaped holes based on generalized regression neural network

Inadequate design of film cooling holes can significantly reduce film cooling effectiveness of turbine blades, compromising high temperature resistance. By comparing experimental data, the most suitable turbulence model is selected to calculate film cooling effectiveness on the leading edge of turbine blades. A high-precision nonlinear mapping relationship between the geometric parameters and arrangement of fan-shaped holes, and film cooling effectiveness is established by using machine learning which is trained by 1200 cases. The genetic algorithm is employed to search for optimal parameters. Sensitivity and correlation analyses are conducted to assess the influence degree of the geometric parameters and arrangement on film cooling effectiveness and these mechanisms are analyzed in depth. Findings reveal that the injection angle exerts the most significant effect on film cooling effectiveness, with significant effects observed for the forward expansion angle and the angle with respect to stagnation line. Interactions among parameters play a crucial role in influencing film cooling effectiveness and cannot be disregarded. In the design space, the optimal configuration comprises the injection angle of 30°, the forward expansion angle of 10°, the streamwise torsion angle of 10°, the pitch of holes of 12 mm and the angle with respect to stagnation line of 30°.

Research on flow and heat transfer characteristics of supersonic film cooling with different hole-configurations

In order to investigates the flow and heat transfer characteristics under supersonic mainstream conditions among three typical film hole-configurations—cylinder hole (CH), fan-shaped hole (FSH), and laidback fan-shaped hole (LFSH), this study conducted numerical simulation using the ANSYS CFX and based on the assumption of compressible flow. The results indicate that the cooling effectiveness of the CH dramatically decreases as the blowing ratio increases, while FSH and LFSH improves. When the supersonic mainstream encounters the coolant, flow structures such as oblique shock waves, expansion waves, and mixing layers et al., will be generated around the film holes, which have an impact on the flow and heat transfer characteristics. Under the conditions of same blowing ratio, the shock wave intensity of LFSH is significantly lower than the other two holes and the kidney vortices formed by the CH develop further along the mainstream direction. While the strength of the kidney vortices formed by the FSH and the LFSH is weaker. Additionally, the LFSH incurred the least total pressure loss from before the shock wave to after the oblique shock wave. Specifically, at a blowing ratio of 0.8, it achieved a reduction of approximately 20% in total pressure loss compared to the CH.

Film-Cooling Effectiveness on a Turbine Blade Platform With Various Hole Shapes and Layouts

Abstract This paper reports a film-cooling effectiveness experiment on a turbine blade platform that combines two film hole shapes with three layouts. A linear cascade using the pressure-sensitive paint (PSP) technique was employed to measure the adiabatic film effectiveness and discharge coefficients. Three film hole layouts, including two double-row layouts and one dispersed layout, were designed based on the platform configurations. One double-row layout arranges both rows of holes on the pressure side. Another double-row layout arranges one row on the pressure side and one row on the suction side. The dispersed layout was designed with streamwise multirows using the same number of holes. A fan-shaped hole and a diffusion slot hole were tested and compared. The experiments were conducted at a mainstream Reynolds number of 7 × 105, a mainstream turbulence intensity of 3.6%, and a coolant-to-mainstream density ratio of 1.5. The blowing ratio ranged from 0.5 to 2.5. The results demonstrated that regardless of the hole shape, the dispersed layout performed better than the two double-row layouts. However, the effects of the layout on the film effectiveness and discharge coefficients are smaller for the diffusion slot hole. In the three layouts, the film effectiveness of the diffusion slot holes is remarkably greater than that of the fan-shaped holes, and the superiority increases as the blowing ratio increases. In contrast, the superiority of the diffusion slot hole in double-row layouts surpasses that of a dispersed layout.

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.

Film cooling performances of short fan-shaped-holes under oscillating freestream

Freestream oscillation presents a significant challenge in optimized design of short fan-shaped-hole, a good choice for more promising double-wall cooled blades. An experimental investigation was conducted to understand detailed effects of key geometrical parameters of the fan-shaped hole, freestream oscillating frequency and cooling air-to-mainstream blowing ratio on the unstable film cooling performances. The selected geometrical parameters included the length-to-diameter ratio, the lateral diffusion angle, and the length ratio of cylindrical section-to-diffusion section. Time-resolved planar quantitative light sheet technique was employed to visualize the temporal variations of jet trajectory and transported scalar concentrations. The experimental results indicated that freestream oscillation causes variations in jet mechanisms, changing the trend in film cooling with blowing ratio and reversing the universally-acknowledged harmful influence of non-fully development of cooling air in short tube. The optimized short-hole can achieve an increment up to 40% in film effectiveness under oscillating freestream, in comparison with the long-hole-jet. The primary principle of the optimized design of short shaped-hole is properly enlarging the lateral expansion angle, aiming at the higher steady film effectiveness while the lower unsteadiness due to the transient fluctuations. Further enlarging the length ratio can improve the stability of film cooling in an oscillating cycle.

Assessment of the effect of thermal barrier coating blockage on film cooling effectiveness: An experimental study of full-coverage turbine vane

The improved performance of turbine blades leads to increased temperature loads, thereby necessitating higher design standards for heat protection. Thermal barrier coating (TBCs) and film cooling are the predominant technologies employed for blade cooling. However, the application of thermal barrier coatings through spraying inevitably modifies the structure of film holes, resulting in deviations from the original design in film cooling. In this study, numerical calculations are used to examine the impact of TBCs blockage on coolant mixing and flow mechanisms in the film holes at typical locations on the vane surface. Additionally, models with actual engine vane dimensions were developed, and the full-coverage film cooling effectiveness (.) vanes with and without TBCs were measured using Pressure-sensitive paint (PSP). The results indicate that the impact of TBCs blockage on the cooling characteristics of cylindrical holes is more pronounced than that of laid-back fan-shaped holes. The TBCs blockage leads to a redistribution of coolant, causing a greater tendency for the coolant to flow out from the hole rows on the suction surface in the tested vane. This results in higher η on the suction surface for the vane with TBCs compared to the vane without TBCs, with a maximum increase of approximately 16.7%. For the other regions of the vane, TBCs spraying significantly reduced η at the leading edge of the vane (with a maximum reduction of approximately 52%) and slightly reduced η at the pressure surface for other regions of the vane (with a maximum reduction of approximately 7%). The application of TBCs reduces the uniformity of film coverage on vane surfaces, with a maximum reduction of approximately 22.5%.

Effect of the hole configurations on effusion cooling effectiveness under swirl impact in gas turbine combustor

Effusion cooling characteristics of the cylindrical and fan-shaped hole configurations are studied under realistic swirl flows at blowing ratios ranging from 1.2 to 6.0. RANS computations with the k-ω SST model are used to evaluate the interaction between swirl mainstream and cooling air. The results show that the cooling effectiveness distribution for the cylindrical and fan-shaped hole configurations are similarly controlled by swirl impact. Two high-temperature regions emerge near the impact location of the swirl main flow on the liner wall. The fan-shaped hole configuration has higher cooling effectiveness, and the difference is relative to location. Quantitatively analyzing, the fan-shaped holes are 19.7 %–53.2 % higher than the cylindrical holes in impact zones. In the corner recirculation zone, the difference ranges from 39.1 % to 84.2 %. The computations reflect the interaction between swirl flows and cooling jets is stronger for fan-shaped holes due to lower outlet velocity. Therefore the cooling air is easier to be suppressed by swirl impact under low BR, while the increasing blowing ratio can enhance the resistance of cooling air against swirl flows.

Experimental optimization of butterfly-shaped film cooling hole configuration with a 30 degree injection angle

Film cooling is one of the external cooling techniques that is widely applied to the hot-gas path components of a gas turbine, forming a thin film of a coolant on the surface to prevent the direct convective heat transfer between combustion gas and components. The film cooling effectiveness (FCE) is significantly influenced by the geometry of the injection hole. Therefore, it is important to adjust the shape parameters in order to derive the optimal hole configuration under given flow conditions. In a previous study, a butterfly-shaped film cooling hole configuration (BSHC), which combines the advantages of fan-shaped holes and anti-vortex holes, was proposed. In this study, an investigation into the optimal shape design of the BSHC was performed using the design of experiments. The forward expansion angle, lateral expansion angle, and twisted angle were defined as design variables. Before the optimization, the effect of each design variable on local and averaged FCE was analysed. Generally, a high FCE could be obtained when the forward expansion angle was small within a range of 7⁰ - 15⁰. However, at high blowing ratios, high effectiveness was also observed even with a large forward expansion angle. Both the lateral expansion angle and the twisted angle had high FCE near the centre point, 11⁰ and 20⁰, respectively. The response surface of the overall averaged FCE was predicted using the Kriging model. The optimized BSHC under a density ratio of 2.0 and blowing ratio of 2.0 was determined and named OPT BSHC. The OPT BSHC showed 8.6 % and 10.5 % higher FCE respectively, compared to two types of the optimized fan-shaped holes that were optimized under the same conditions.

Experimental and numerical investigations on film cooling characteristics and hole shape improvement of turbine vane at one engine inoperative condition

This study investigates the film cooling characteristics of four cooling schemes of a gas turbine vane at various mass flux ratios (8.0%, 8.5% and 9.0%) using the Pressure Sensitive Paint (PSP) technology. The improved cooling scheme involves the alteration from cylindrical holes to laid-back fan-shaped holes and laid-back holes on the pressure surface and the leading edge. Considering the complex fluctuations of the real engine operating environment, the one engine inoperative (OEI) situation is numerically simulated to explore the turbine vane's cooling performance at varying operating conditions. Experimental and numerical results indicate significantly enhanced film cooling effectiveness of vane with shaped holes (improved cooling schemes) compared to the cylindrical hole (baseline), particularly at higher mass flux ratio (MFR = 9.0%). At three OEI conditions, the film coverage of improved cooling schemes is significant compared to the baseline on the pressure side. When the OEI condition is increased from 100% to 120%, the area-averaged cooling effectiveness of the shaped hole vane decreases by 4.3%, due to the interaction between the freestream and lift-off cooling jet.

Investigation of the effect of combustor swirl flow on turbine vane full coverage film cooling

In response to the challenges posed by aviation engine energy consumption and the imperative to reduce NOx emissions, modern advanced aviation engines have increasingly embraced lean-burn and swirl-stabilized combustion chambers. Consequently, the impact of residual swirl flow at the combustion chamber exit on heat transfer to turbine blades has gained prominence. This study delves into the cooling characteristics and variations under swirling inlet conditions, employing both pressure-sensitive paint (PSP) experiments and numerical simulations. The experiments measured film cooling effectiveness (η) on full film cooling blades across different levels of swirl inflow, while numerical simulations dissected flow patterns under swirling and uniform inflow conditions. The results reveal that the extent of film migration does not exhibit a direct proportionality to the level of swirl intensity. Under low swirl conditions (SN = 0.3), film migration is subtle, resulting in a maximum area-averaged η reduction of 3.0%. Conversely, under high swirl conditions (SN = 0.45), film migration becomes notably pronounced, leading to a maximum area-averaged η reduction of 11.9%. Increasing the coolant flow rate mitigates the migration characteristics, albeit with limited effectiveness. The migration of film on the blade surface of full film cooling blades is partially induced by swirling vortices but is primarily attributed to the migration of the leading edge film. Swirling inflow induces the movement of the stagnation-point line, thereby affecting the distribution of cooling air on the blade surface. Finally, this study proposes a combination of measures as an improvement strategy, including an asymmetric counterflow arrangement for the leading-edge holes, a reduction in the outflow angle of the leading-edge holes, and forward-laidback fan-shaped holes. These improvement measures are anticipated to alleviate film migration, reduce the overall uneven film distribution, and increase the average η across the blade.

Effect of the Key Geometry and Flow Parameters on Discharge Coefficient of Laidback Fan-Shaped Hole Under Coolant Crossflow Condition

Abstract A laidback fan-shaped hole is commonly used due to its superior lateral film coverage. Its discharge coefficient is significantly influenced by internal crossflow owing to its complex geometrical structure. In this paper, the authors numerically investigate the flow mechanisms of the laidback fan-shaped hole under the influence of internal crossflow. The numerical simulations utilize the validated SST k–ω turbulence model, with the Reynolds number of internal crossflow ranging from 20,000 to 160,000 and the ratio of pressure ranging from 1 to 1.6. The results show that the different orientations of internal crossflow cause varying degrees of in-hole separation that led to a discrepancy in the discharge coefficient. The larger the Reynolds number of the crossflow is, the more drastic the change in the discharge coefficient. Furthermore, a comparison between the results obtained with and without internal crossflow has shown that the length of the cylindrical section is the primary factor determining the discharge coefficient of the laidback fan-shaped hole. The magnitude of the discharge coefficient depended on the extent of flow separation within the cylindrical section. Additionally, the numerical simulations obtained the discharge coefficient under a high internal crossflow Reynolds number of internal crossflow and a wall with a constant thickness and compared it with the predictions of a low-dimensional model of the discharge coefficient (based on our previous experimental data). The discrepancy between the results is within 10%, thus verifying the scalability of the low-dimensional model.

Non-axisymmetric Endwall film cooling characteristics considering the influences of cylindrical holes and laidback fan-shaped holes

Flow fields near the turbine vane endwall are complicated due to the endwall cross flows. The use of a non-axisymmetric endwall is regarded as an efficient technique to reduce the lateral pressure difference, decreasing the endwall cross flow. Numerical analysis was performed to determine how the non-axisymmetric endwall affected the vortex structure and heat transfer level. The cooling performance was investigated with cylindrical and laidback fan-shaped holes (7–7–7), which were arranged in rows aligned in the axial direction. The results showed that the non-axisymmetric endwall could significantly reduce the circumferential pressure difference and suppress the growth of the passage vortex, and the area-averaged heat transfer coefficient was reduced by 3.34%. The outlet area of the film hole was altered by the non-axisymmetric endwall, and the over-cooled regions may have appeared as a result of the excessive area increase. The influence of the non-axisymmetric endwall was concentrated at 0.4 < Z/Cax < 1.0 for the cylindrical hole. With the increase in M, the film cooling effectiveness of the non-axisymmetric endwall attained a higher level than that of the flat endwall. For the laidback fan-shaped hole, the effect of the non-axisymmetric endwall was confined within 0.25 < Z/Cax < 1.0. The half-period trigonometric function of the non-axisymmetric endwall (HTFN) achieved the optimal cooling performance for three blowing ratios. However, the periodic trigonometric function of the non-axisymmetric endwall (PTFN) only outperformed the flat endwall when M= 1.5.

Film cooling and stress concentration properties of the backward-diffusion elliptical hole

Film cooling is a commonly used cooling technology in gas turbines. However, it faces the problem of stress concentration. The present work proposes a low-stress film cooling hole with good cooling performance, namely the backward-diffusion elliptical hole. The Pressure Sensitive Paint experiments compare the film effectiveness of the backward-diffusion elliptical hole and the 7–7–7 fan-shaped hole. The film effectiveness of the backward-diffusion elliptical hole is higher than that of the 7–7–7 fan-shaped hole, especially at high blowing ratios. The structure of the jet vortex system of the backward-diffusion elliptical hole is analyzed by large eddy simulation. The results show that the backward-diffusion elliptical hole induces a stronger anti-counter-rotating vortex pair than the elliptical hole. The anti-counter-rotating vortex pair weakens the harmful effects of the counter-rotating vortex pair on the film effectiveness. The stress concentration factor of the backward-diffusion elliptical hole is obtained by finite element analysis. The backward-diffusion elliptical hole’s stress concentration factor is about 2.85, similar to the elliptical hole (2.84). However, the stress concentration factor of the 7–7–7 fan-shaped hole is 6.58. Therefore, the stress concentration factor for the backward-diffusion elliptical hole is only 43.3% of that for the 7–7–7 fan-shaped hole. The backward-diffusion elliptical hole is a more promising low-stress, high-cooling-efficiency film hole.

TPIV Experimental Investigation of Film Coolant-to-Mainstream Interaction From Shaped Cooling Holes With Various Inlet Geometries

Abstract In this investigation, multiple sets of time-averaged tomographic particle imaging velocimetry measurements are completed for laid-back, fan-shaped film cooling holes with “racetrack” shaped inlets. Traditional 10-10-10, laid-back, fan-shaped holes are inclined 30 deg to the mainstream flow on a flat plate. The inlet cross section varies from round to two elongated racetrack shapes. The cross-sectional area and the outlet-to-inlet area ratio for all the geometries are held constant. The flat plate is installed in a low-speed wind tunnel with a mainstream turbulence intensity of 8% and an average velocity of 21.6 m/s. The blowing ratios of the film jets range from 0.6 to 1.5 and the density ratio is 1. The Reynolds number of the cooling jet varies from 2600 to 8400. The characteristics of the resulting flowfield are coupled with the detailed film cooling effectiveness distributions. It can be noted from the results that the counter-rotating vortex pair generated by the 2:1 inlet is the closest to the surface and weakest in strength, likely caused by the minimum peak jet momentum of the three. The Reynolds stresses downstream of the 2:1 and 4:1 inlets are significantly lower than those downstream of the shaped hole with a round inlet. An inverse relation between volumetric turbulence accumulation (TA), and surface effectiveness (η), can be correlated for the blowing ratios considered. The turbulence accumulation term can thus be used to evaluate the performance of a film cooling hole design with flowfield data.