Film Cooling Performance Research Articles

Novel Shaped Sweeping Jet for Improved Film Cooling and Anti-Deposition Performance

Novel Shaped Sweeping Jet for Improved Film Cooling and Anti-Deposition Performance

The Effect of Reverse and Opposite Injection on the Film Cooling Performance of a Symmetrical Turbine Blade

The Effect of Reverse and Opposite Injection on the Film Cooling Performance of a Symmetrical Turbine Blade

A Comprehensive Investigation of the Effect of Hole Diameter and Number on Film Cooling Performance of Rotating Blade Leading Edge

The combined effect of film hole number and diameter on the film cooling were investigated on the rotating blade leading edge by using experimental and numerical methods. The film cooling performance was obtained by pressure sensitive paint technique under rotational condition. To better understand the experimental results, the flow fields obtained from numerical simulation were also analyzed. The effects of hole length-to-diameter ratio (L/d=2.5/5/7.5), hole number (N=7/9/10/11/13), hole diameter(d=0.64∼1.26mm), Hole to hole pitch, (P/d=6.3∼9.2) and hole exit area (Ahole/ALE=4%-14%) on the film cooling performance are investigated. Five blowing ratios are also employed (M=0.5 to 2.0). The results indicate that the hole length-to-diameter ratios (L/d=2.5∼7.5) have a negligible influence on the film cooling effectiveness. Both the film hole number (N=10/13) and the hole diameter (d=0.64/0.75mm) monotonically affect the film cooling performance under each blowing ratios. However, a comprehensive study indicates that there is a significant quadratic correlation between the area-averaged cooling effectiveness and the total film hole exit area for different combinations of hole number and diameter. Meanwhile, there is an optimal film hole exit area on the leading edge and it increases as the blowing ratio or mass flow rate increases.

Coupling Effect of Particle Deposition Inside and Outside Holes on Film Cooling Performance on the Leading Edge of the Blade

Coupling Effect of Particle Deposition Inside and Outside Holes on Film Cooling Performance on the Leading Edge of the Blade

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.

The Evolution of Flow Structures and Coolant Coverage in Double-Row Film Cooling with Upstream Forward Jets and Downstream Backward Jets

The spatiotemporal evolution of the flow structures and coolant coverage of double-row film cooling with upstream forward jets and downstream backward jets, having a significant impact on film-cooling performance, is studied using the simplified thermal lattice Boltzmann method (STLBM). Moreover, the effect of the inclination angle of downstream backward jets is considered. The high-performance simulations of film cooling have been conducted by using our verified in-house solver. Results show that special flow structures, such as a sand dune-shaped protrusion, appear in double-row film cooling with upstream forward jets and downstream backward jets, which is mainly because of the blockage effect resulting from the coolant jet with backward injection. The interaction among structures results in the generation of an anti-counterrotating vortex pair (anti-CVP). The anti-CVP with the downwash motion can result in the attachment of coolant to the bottom wall, which promotes the stability and lateral coverage of coolant film. The momentum and heat transport are strengthened as the backward jet is injected into the boundary layer of the mainstream. Although the downstream evolution of the backward jet is not very smooth, its core attaches closely to the bottom wall due to the downwash motion of anti-CVP. Moreover, there is an obvious backflow zone shown in the trailing edge of the downstream backward jet with a large inclination angle. The obvious backflow makes the coolant attach to the bottom wall well. Therefore, the film cooling effectiveness is improved as the inclination angle of the downstream backward jet varies from αdown=135° to αdown=155°, with a constant blowing ratio of BR=0.5. In addition, the fluctuation of the bottom wall’s temperature is weak due to the stable coverage of the coolant layer under αdown=155°. The film-cooling performance with an inclination angle of αdown=155° is the best among all the cases studied in this work. This work provides essential insights into film cooling with backward coolant injection and contributes to obtaining a complete understanding of film cooling with backward coolant injection.

Film Cooling Performances Under Various Upstream Roughness Conditions: Experimental Investigations and Similarity Hypothesis

Abstract Roughness caused by deposition, erosion, and additive manufacturing can significantly affect gas turbine efficiency. Previous research has often examined film cooling performance under limited roughness configurations, resulting in inconclusive findings. In this study, film cooling performances under various upstream roughness conditions were investigated to simulate the roughness-affected film cooling performance of the suction side and the endwall. Three roughness heights (k/D = 0.1, 0.2, and 0.4) and shapes (ks/k = 0.17, 0.67, and 1.95) were selected to cover a wide range of surface roughness characteristics. Three blowing ratios were examined (M = 0.5, 1.0, and 1.5). The weakly roughened surfaces showed improved cooling effectiveness as k/D increased. Meanwhile, the moderately and severely roughened surfaces showed a decrease in cooling effectiveness with increasing k/D at M = 0.5 and 1.0 but an increase at M = 1.5. Cases with shallower and higher roughness elements at M = 1.5 outperformed the smooth plate. Subsequently, a similarity hypothesis for film cooling effectiveness was proposed. At all blowing ratios, the scaled cooling effectiveness profiles converged around the smooth plate results for ks/D < 0.391, encompassing common turbine roughness scales, including irregularly roughened surfaces. Deviations emerged at ks/D = 0.782, and they were correlated with the deteriorated regions observed at various blowing ratios. Ensemble-averaged scaled cooling effectiveness exponentially grew with increasing roughness scale for all blowing ratios, and empirical expressions based on smooth plate result and roughness scale were proposed.

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.

Heat Transfer Coefficient and Adiabatic Effectiveness Measurements on a Nozzle Guide Vane With a Single Row of Cylindrical Holes

Abstract Film-cooling injection significantly affects the thermal behavior of turbine vane surfaces. In addition to the beneficial effect of the film shielding the vane from the hot gas flow, alteration of the thermal boundary layer should also be taken into account. The aim of the present work is to detail the film-cooling performance in terms of adiabatic effectiveness and external heat transfer coefficient on a 2D nozzle guide vane. A single row of cylindrical holes was tested on both pressure and suction sides of a literature vane, the VKI LS89, in a linear cascade. The employed measurement technique is a transient thermal method based on infrared thermography, which was thoroughly described and validated in a previous work. The influence of inlet freestream turbulence and blowing ratio was evaluated, and two different injection angles were considered for both pressure and suction sides. Spatially resolved distributions of adiabatic effectiveness and heat transfer coefficient (HTC) on the vane surface allow us to precisely quantify the above-mentioned aspects and highlight qualitative differences between pressure side and suction side behavior. Details regarding the generated non-uniformities in the measured parameters could be also provided, to emphasize how average quantities are not always sufficient to characterize such complex phenomena. The impact of different reference conditions to scale HTC results was also investigated. Such effect was found not negligible on the overall performance of the film-cooling system, especially on the suction side where transition plays a critical role. Ultimately, the collected results constitute a wide and detailed experimental database for numerical modeling validation in a well-studied environment as the LS89 configuration.

Effects of Downstream Vortex Generators on Film Cooling a Flat Plate Fed by Crossflow

Abstract Counter-rotating vortices, formed by the interaction of film-cooling jets and the hot gas flow, adversely affect the performance of conventional film-cooling designs. Downstream vortex generators have been shown to improve cooling effectiveness by mitigating the effects of the counter-rotating vortices and by deflecting the cooling jet laterally. In this study, computational and experimental methods were used to examine how cylindrical film-cooling holes (D = 3.2 mm, L/D = 6, p/D = 3, α = 30 deg) with and without downstream vortex generators perform when the coolant supply channel is perpendicular to the direction of the hot gas. For this study, the hot gas had a temperature of 650 K and an average Mach number of 0.23. The hot-gas-to-coolant temperature ratio was 1.9, and two blowing ratios (0.75 and 1.0) were studied. Results from the computational fluid dynamics study show how crossflow affects the interaction between the film-cooling jet and hot gas flow with and without downstream vortex generators. The experimental measurements were based on infrared thermography in a conjugate heat transfer environment. Results were obtained for film-cooling performance in terms of overall effectiveness, film effectiveness, and local heat transfer coefficients. The downstream vortex generators can increase the laterally averaged effectiveness by a factor of 1.5 relative to cylindrical holes, but this higher performance is restricted to low crossflow velocities and higher blowing ratios.

Effect of Plasma Actuators on Film Cooling Performance with Forward Injection Slots: A Numerical Approach

Effect of Plasma Actuators on Film Cooling Performance with Forward Injection Slots: A Numerical Approach

Turbine Blade Endwall Heat Transfer and Film Cooling Performance With Multigap Jets and Film Holes

Abstract Due to manufacturing, assembly, and operational conditions, gaps are commonly presented in the turbine components, such as platform blade-to-blade gap (slashface gap) and rim cavity platform gap (slot gap). The coolant ejected through these gaps not only limits hot mainstream gas ingestion into the disk cavity but also has the potential to provide endwall film cooling coverage. Nevertheless, the interaction of these gap flows, predesigned film hole jets, and endwall secondary flows, significantly impacts endwall heat transfer and film cooling performance and leads to a more complicated flow field near endwall. To further understand the endwall flow physics behavior in this complicated flow field, the combined effects of film hole jets and multigap leakages (slashface jet and slot jet) were experimentally studied in a transient wind tunnel with a six-blade linear cascade (nonrotating). The endwall heat transfer was measured and recorded by the IR technique at inlet average turbulence intensity (Tu) of 7.5% and an exit Mach number (Maex) of 0.4. In addition, detailed numerical predictions were also performed to discuss the flow physics near endwall and the multigap leakages behavior. Results indicated that the slashface jet is an essential contributor to endwall film cooling performance in this endwall configuration. The slashface jet delays the development of the film hole jets, leading to an enhancement of film cooling performance downstream of the film holes. Increasing the mass flow ratio of the slashface jet (MFRslashface) can prevent the high-temperature mainstream ingestion, reduce the thermal failure risks, and increase the peak of film cooling effectiveness downstream of the endwall (167% at MFRslashface = 0.5%). Due to the limitation and hindrance impacts of the disk vortex (DV), the cavity vortex (CV), and the suction side leg of the horse-shoe vortex (HV,s), the slot flow is confined to a small region downstream of the slot gap.

Large eddy simulation of film cooling for a forward expansion hole under main flow oscillation

This study conducted a large eddy simulation (LES) to predict the flow structure and effectiveness of film cooling in a gas turbine. The study focused on a forward expansion hole on a flat plate with an inclined injection angle of 35° and spanwise injection angles (orientation angles) of 0° and 30° under main flow oscillation. In gas turbines, the mainstream flow can exhibit unsteadiness owing to stator–rotor interactions or freestream turbulence. Here, sinusoidal functions are used to approximate the flow oscillations, and their effects are evaluated at frequencies of 0, 8, and 32 Hz, corresponding to Strouhal numbers of 0, 0.603, and 2.413, respectively. The study aims to characterize the film cooling at two time-averaged blowing ratios: 0.5 and 1.0. The range of Strouhal numbers considered in this study is relevant to actual transonic gas turbines. The computational results obtained through LES are validated using experimental data and compared with the Reynolds-averaged Navier–Stokes simulation results. The findings demonstrate the efficiency of LES to predict accurately the distribution of spanwise film cooling effectiveness and spanwise-averaged effectiveness for forward expansion holes on turbine blades under flow oscillations. Comprehensive analyses of various aspects such as the instantaneous turbulent structures of the film cooling flow fields, time-averaged temperature, and root-mean-squared values of the coolant temperature and a proper orthogonal decomposition analysis for velocity vectors are performed. The results show that when the oscillation frequency of the main flow varied from 0 Hz to 32 Hz for a time-averaged blowing ratio of 0.5, the film cooling effectiveness for the forward expansion hole reduced significantly. However, when the frequency increased under a time-averaged blowing ratio of 1.0, the film cooling effectiveness for the hole was only slightly affected by the oscillations. Moreover, an orientation angle of 30° enhanced the film-cooling performance even under flow oscillations. Thus, this paper reports the development of highly effective film-cooling strategies that can guide researchers aiming to mitigate the adverse effects of flow oscillations, thereby enhancing the reliability and efficiency of gas turbine engines.

Large Eddy simulation in optimizing trench configuration to improve cooling effectiveness of a laidback fan-shaped hole

In modern gas turbines, a thermal barrier coating (TBC) is used to protect the components of the turbine from the hot flow coming from the combustor, which results in the cooling holes on the blade surfaces being embedded in a 2D trench slot. In this study, large eddy simulations (LES) were carried out to investigate the effects of trench geometry on the film-cooling effectiveness of a shaped cooling hole. The shaped cooling hole is located on a flat surface with an injection angle of 30 degrees, and the blowing and density ratios are fixed at 1.5. Thirteen different designed cases were sampled using Box-Behnken method based on three different geometry parameters of the trench slot, trench height and upstream and downstream trench lengths. A genetic aggregation algorithm was applied to achieve the optimal trench configuration that maximizes the overall cooling effectiveness. The film-cooling performance computed by the LES method was validated for a reference case without a trench and three different trenched cases with the experiments using the PSP (pressure sensitive paint) technique. The cooling effectiveness of the trench cavity and flat plate was found to be significantly enhanced in cases of trenches with relatively greater trench height and less downstream trench length. The optimal trench configuration demonstrated less turbulent fluctuations in the mixing region, and the overall area-averaged adiabatic film-cooling effectiveness was improved by 23% in comparison with a reference hole without a trench.

Cooling performance of film-cooling holes fed by channels of various shapes

Internal and external cooling is pertinent to maintaining operable surface temperatures of hot-section gas turbine components. Film cooling on the external surfaces is supplied from cooling passages placed inside of the blades and can affect the external film-cooling performance. Assessing internal cooling designs that were once difficult to investigate because of the complexity and costs of manufacturing cast turbine blades is now possible to more quickly and less costly through additive manufacturing (AM).This paper provides results and analysis of an experimental and computational study of the cooling performance of meter-diffuser shaped film-cooling holes fed by a range of individual channel shapes with different cross-sectional shapes: circle, hexagon, pentagon, ellipse, diamond, square, rectangle, trapezoid, triangle – vertex up, and triangle – vertex down. These geometries were made possible through the use of AM. The perimeter, and consequently the surface area of the supply channels to the cooling holes, was maintained constant between the different channel shapes. The predictive computational results showed that the shape of the channel affects the secondary flows through the hole that in turn affects the overall film cooling as determined experimentally. The triangle – vertex up created flow structures in the diffuser-shaped cooling hole that caused the coolant jet to create a film on the surface with wide lateral spread; whereas the triangle – vertex down created flow through the cooling hole that caused the jet to fully detach at the same blowing ratio. Consequently, the triangle – vertex up had a 14% improvement in overall effectiveness over the triangle – vertex down and a 33% improvement over the circle which had the lowest overall effectiveness of all the tested shapes.

Reynolds stresses and turbulent heat fluxes in fan-shaped and cylindrical film cooling holes

Large eddy simulation (LES) is carried out to predict the thermal performance, unsteady flow patterns, and turbulence characteristics of the fan-shaped hole and the cylindrical hole cooling films. Velocity fields, Reynolds stresses, and turbulent heat flux distributions are the factors that ultimately determine the film cooling effectiveness by deciding the concentration of coolant and heat near the surface. Accurate prediction of these processes can provide a reliable data set for the design of film cooling systems. It can also help solve difficulties of improving RANS-based simulations in terms of their inaccuracies in matching measured flow rates for a given pressure drop across the film and the mixing process once the coolant is ejected from the film. Credible estimates and modest costs of the current LES originate from the multiblock hexahedra meshing, the turbulence inflow generator, and the numerical schemes implemented. Multiblock hexahedra meshes are adopted and generated by a quasi-automatic mesh generation algorithm utilizing template-based multi-block construction followed by elliptic smoothing. Computational resources are utilized in a way that mesh refinement exists only near cooling decorations of small size and wall boundaries, without the need for associated clustering in the far field which is commonly used in commercial packages. The high orthogonality and continuous smoothness throughout the computational domain are maintained to facilitate turbulence capturing. The synthetic inflow generator produces a random field matching a realistic set of two-point statistics based on eigen-reconstruction in a computationally inexpensive processing step. To prove that the current LES simulations of modest cost can reproduce with high accuracy, the discharge characteristics, velocity, and turbulence intensity profiles of cylindrical and fan-shaped cooling films are compared with published measured data on the same flow configurations. Results reveal that the set of modeling methodologies is valid for calculating film-cooling performance within an acceptable range of accuracy. The energy-carrying turbulence structures in and near the cooling holes are shown and their behaviors are linked to turbulent Reynolds stresses. Reynolds stresses and turbulent heat fluxes are compared for different blowing ratios and two types of cooling holes in the parametric study. Similarity patterns of velocity profiles, non-dimensional temperature profiles, Reynolds stress and turbulent heat flux profiles are discussed.

Experimental and Numerical Study of the Effect of Double Row Slot Injection Locations on Film Cooling Performance of a Corrugated Surface

ABSTRACT Based on experimental and numerical studies the present work proposes an effective injection configuration for film cooling applications on sinusoidal corrugated surfaces. The experimental study is performed for different double-row injection locations (per wavelength) viz. (L0-L25), (L0-L50), (L0-L75), (L25-L50), (L25-L75) and (L50-L75) on the sinusoidal corrugated surface at blowing ratio (1), fixed injection angle (45°), and density ratio (1.095). The numerical study is conducted for a wide range of operating parameters such as blowing ratios on (0.5–2), and density ratios (1.095, 1.5, and 2.5). To access the effect of corrugation geometry three different amplitude to wavelength viz. 0.05, 0.075, and 1 are considered in the present numerical study. The outcome of the present double-row slot injection investigations on the sinusoidal corrugated surface indicates that the flow pattern and film cooling strongly depend on secondary injection locations. The (L50-L75) injection case shows the best performance among all selected cases. It shows a 15–19% improvement in film cooling effectiveness (average over one wavelength). The increase of amplitude to wavelength ratio is reported to be a detrimental effect and decreases the film cooling effectiveness. The increase of corrugation amplitude to wavelength ratio increases the depth of the corrugation valley, further restricts the coolant spread, and promotes coolant mixing with hot mainstream. The coolant injection in mainstream flow was also found to be a distinct impact on flow and the nondimensional velocity profile. Apart from that a significant impact of the blowing ratio and density ratio is reported.

Effects of Crescent Shaped Block Width and Length on Flow Field and Film Cooling Performance

Turbine blades and rocket nozzles can be efficiently protected from combustion chamber exhaust hot gases by film cooling technic. A crescent-shaped block placed downstream of the injection hole can significantly improve cooling effectiveness. The main objective of this research work is to investigate the influence of the crescent-shaped block length and width on film cooling effectiveness. ANSYS CFX was used to conduct analysis of a flat plate configuration with cylindrical holes. Nine configurations of the crescent shaped blocks were considered. For each case, the effect of blowing ratios (0.5 and 1) was investigated. The turbulence model shear stress transport (SST) is used for approximating turbulence. Good agreement was obtained by comparing the analysis results with the experimental data. The result indicated that block width variation has a considerable impact on film cooling performance. However, slight effect of block length on cooling effectiveness was obtained. Comparing all analyzed configurations, the best cooling effectiveness was reached for the model 9 (W = 3d, B = 1.5d). HIGHLIGHTS Rocket nozzles can be efficiently protected from combustion chamber exhaust hot gases by film cooling technic One of the latest techniques to optimize the overall performance of film cooling is adding an obstacle configuration downstream the coolant perforation Film cooling performance is considerably enhanced by placing a crescent shaped block downstream the injection hole The maximum enhancement in overall film cooling effectiveness using the optimal configuration (model 6: W = 2d, B = 1.5d) is about 252 % achieved at a blowing ratio of M = 1 GRAPHICAL ABSTRACT

Effect of Slot Jet Flow on Non-Axisymmetric Endwall Cooling Performance of High-Load Turbines

As vane inlet temperatures and turbine loadings are increasing, the aerodynamic and thermal management for the endwalls of gas turbines have received increased attention. Non-axisymmetric endwalls are becoming popular due to their proficient capabilities to modify the secondary flow fields and to change the film cooling performance on the endwalls. In this study, by considering the interaction between mainstream and purge flow based on non-axisymmetric endwall contouring, the numerical research model used in the present research was established. Based on the validated numerical method, the influence of the non-axisymmetric endwall contouring on the film-cooling effectiveness and aerodynamic characteristics was studied. Furthermore, the effect of different inclination angles on the film-cooling performance of the contoured endwalls was also investigated. The results indicate that for the high-load turbine vane used in this research, various types of non-axisymmetric endwall contouring can alter the aero-dynamic characteristics and cooling performance simultaneously. By inhibiting the secondary flows, non-axisymmetric endwall contouring can reduce the total cascade pressure loss coefficient by 0.305%. In addition, non-axisymmetric endwall contouring can significantly enhance the effective coverage area of purge flow up to 28.29%, and the endwall near the suction side can achieve better cooling performance. Finally, non-axisymmetric endwall contouring can improve the protective effect of large-angle purge flow.

Film Cooling Performance of a Cylindrical Hole with an Upstream Crescent-Shaped Block in Linear Cascade

Recent works have already demonstrated that placing a crescent-shaped block upstream of a cylindrical hole could enhance the cooling performance of flat-plate films. The flow and cooling performance of the crescent-shaped block applied over the pressure and suction sides of the blade is investigated in this article. The Reynolds-averaged Navier-Stokes equations are solved with the Shear Stress Transport model for turbulence closure. Two optimized blocks are obtained from the flat-plate film cooling in our previous work, and two positions on the pressure and suction sides are tested. The blowing ratio varies from 0.5 to 2.0. The results show that when the block is applied on the blade surface, it yields a different cooling performance compared with the flat plate due to different geometry curvature and pressure gradient. The cooling performance on the suction side is slightly higher than on that on the pressure side, while the aerodynamic loss on the suction side is much higher. For the different blocks, the qualitative change of cooling performance vs. blowing ratios held on turbine blades is quite close to that of flat plates. The optimized smaller block in the flat plate provides better cooling performance at lower blowing ratios, while the larger block is superior when the blowing ratios are higher.