Film Cooling Effectiveness Research Articles

Enhanced cooling performance of blade tip slot cooling: Effect of slot open length

Addressing the critical challenge of thermal load-induced damage in turbine blade tips, this study aims to elucidate the effects of slot open length on film cooling effectiveness (FCE) across turbine blade tips with Pressure Side (PS) and Suction Side (SS) slots considering the application of additive manufacturing to turbine blade tips. Employing a linear cascade setup and the pressure sensitive paint (PSP) technique, we examined and quantified cooling performance across turbine blade tips with PS or SS slots at various slot open lengths. We revealed that while short slot open lengths enhance FCE at the leading edge (LE) due to increased averaged blowing ratios, they fail to adequately cover the trailing edge (TE) region. This decrease in cooling performance is largely attributable to the mixing of coolant with hot gas. In contrast, longer slot open lengths achieve a consistent FCE distribution across the blade tip despite a slight decrease in FCE at the LE region. The study suggests that integrating additional thermal design strategies, such as partition walls with slots, could effectively manage hot gas behaviors, thus enhancing the overall cooling design of turbine blade tips. This research makes a significant contribution to the field of turbine blade tip cooling design, providing insights into optimizing cooling performance through advanced slot configurations enabled by additive manufacturing technologies.

Simulations With Compressible Multiphase Formulation on Heat Transfer Studies in a Rocket Thrust Chamber With Experimental Validation

Abstract In the combustor for rocket engines, liquid film cooling is a widely adopted technique for restricting the structural components to the acceptable bounds for the successful operation of the propulsion systems. Multiple cooling techniques for thrust chamber walls are favored along with the coolant film in propulsion systems operating at higher chamber pressures by incorporating an ablative cooling mechanism for the nozzle. The adoption of combination will result in challenges in simulation studies due to the presence of multiphase phenomenon in the system. The current article presents the effects of these two cooling methods on a thrust chamber wall studied through compressible multiphase formulations. A majority of the previous studies have used incompressible flow equations to model coolant film behavior and associated heat transfer. The present study utilizes an approach utilizing compressible multiphase flow to simulate the behavior of the compressible hot gas flow and the incompressible coolant liquid film accounting for phase change effects, vapor diffusion into hot gas, radiation effects on coolant surface, liquid entrainment, and blowing effects due to vapor formation at interface. Film-cooling effects on ablative nozzle accounting for pyrolysis phenomenon due to material decomposition governed by Arrhenius expression are also emphasized. The outcome of the numerical model showed good agreement with available test data in literature, and results from the tests done in-house. The model described in the present study was able to support the actual evidence of liquid film cooling being able to preserve the walls of the thrust chamber from severe internal thermal environment and prevent ablation on surface of nozzle.

A Numerical Investigation of Particle Swarm Optimization to Improve Film-Cooling Effectiveness from Converging Slot Hole

A Numerical Investigation of Particle Swarm Optimization to Improve Film-Cooling Effectiveness from Converging Slot Hole

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 investigation on the effects of complex pressure gradients on the film cooling effectiveness of a serpentine nozzle

Serpentine nozzles more easily deform and damage due to high-temperature airflow because of their multi-bend channel configuration. Therefore, film cooling technologies should be applied to serpentine nozzles. Complex pressure gradients in serpentine nozzles lead to different film cooling characteristics from combustion chambers, turbine blades, and other exhaust systems. This study experimentally investigated the effects of complex pressure gradients and film hole inclination angles on a serpentine nozzle’s film cooling effectiveness (FCE) using pressure-sensitive paint (PSP). The flow mechanisms were obtained via numerical simulations. The experimental nozzle design needs to consider the needs of PSP observations and the normal flow of serpentine nozzles. However, the complex structure of serpentine nozzles makes those increasingly difficult. The studied pressure gradients include the pressure gradient mutation (PGM), strong adverse pressure gradient (SAPG), weak adverse pressure gradient (WAPG), and favorable pressure gradient (FPG). The results show that complex pressure gradients change the coolant coverage and adhesion and may induce a recirculation zone in the SAPG region. This affects the development of the vortex construction and leads to different FCE distributions. The FCE under the effects of the PGM is the largest, followed by the SAPG, WAPG, and FPG. A larger inclination angle weakens the coolant adhesion but may enhance its downstream and lateral flow treads. This is coupled with the effects of complex pressure gradients, where larger inclination angles bring better film cooling effects in some cases. After accounting for the four blowing ratios, the area-averaged FCEs under the effects of the SAPG, WAPG, and FPG are 28.9 %, 42.8 %, and 52.3 % lower than that under the PGM, respectively. The area-averaged FCEs under the conditions α = 45° and 60° are 6.6 % and 10 % lower than that under α = 30°.

Effect of Inlet Geometry on Film-Cooling Effectiveness from Shaped Holes

Modern gas turbine engines require a sophisticated cooling system design to achieve higher power output and efficiency. This study investigates the potential effect of noncylindrical inlet geometries on the performance of laid-back, fan-shaped film-cooling holes using the pressure-sensitive paint measurement technique on a flat plate. On the basis of a common pattern of outlet geometry, racetrack-shaped inlet geometries with aspect ratios of 2:1 and 4:1 were tested along with traditional cylindrical inlets. The coolant flow conditions range from M = 0.3–1.5 and DR = 1 and 2. The mainstream turbulence intensity is held at 6%. Test results show that the shaped inlets provide a higher area-averaged film-cooling effectiveness over the cylindrical inlet using the same amount of coolant. For the 2:1 inlet, an advantage of 20% higher effectiveness could be maintained for DR = 1, while for DR = 2 this advantage is reduced to 10%. For the 4:1 inlet, when the coolant momentum flux ratio I < 0.5, a similar or slightly higher improvement can be obtained, but when I > 1, the advantage diminishes with the growing I to approximately 5%, at I = 2.25. Regarding discharge coefficients, the 2:1 inlet geometry is similar to the cylindrical inlet. For the 4:1 inlet, it is 2–5% lower or nearly equivalent.

Influence of Relative Position of the Initial Region of Double-Row Perforation of the Turbine Vane Suction Side on the Film Cooling Effectiveness in Conditions of Inlet Flow Irregularity

Influence of Relative Position of the Initial Region of Double-Row Perforation of the Turbine Vane Suction Side on the Film Cooling Effectiveness in Conditions of Inlet Flow Irregularity

Experiments on heat transfer and film cooling characteristics of double-wall cooling structure with novel slot holes and pin-fins

The influences of blowing ratio, mainstream velocity, density ratio and jet Reynolds number on the heat transfer and film cooling properties of double-wall cooling cell (DWCC) with slot holes, impingement and diamond-shaped pin-fins were studied by means of pressure sensitive paint and transient liquid crystal. The simulation obtained the flow characteristic at a jet Reynolds number of 20000. The results indicated that with the increase of blowing ratio, the film cooling performance of DWCC structure also increased, and there was no film lift-off phenomenon which occurred in the cylindrical hole. The presence of pin-fins created a spike-shaped film in the slot outlet. Therefore, when studying the DWCC structure, it was important to consider the couple effects of impingement, pin-fins, and slot hole. As the density ratio decreased and the mainstream velocity increased, the film cooling effectiveness (FCE) gradually decreased. When jet Reynolds number increased from 20,000 to 60,000 under a density ratio of 2.0, the area-averaged FCE increased from 0.308 to 0.435, with an increase of up to 41.2%. The convective heat transfer level of impingement stagnation point was much higher than that of pin-fins region. Due to the transition between the impinging jet and the wall jet, the heat transfer coefficient (HTC) had a multi-peak distribution. In addition, several low heat transfer zones similar to wing-shaped, eye-shaped, and crescent-shaped were observed on the target surface and bottom surface inside the cooling cell. The DWCC structure studied in this paper is more efficient in cooling performance compared to traditional cooling structures. This makes it suitable for use in various fields such as heat transfer of turbine blades and electrical equipment.

Swirling flow field reconstruction and cooling performance analysis based on experimental observations using physics-informed neural networks

The design of thermal protection modules (such as film cooling) for combustion chambers requires a high-fidelity swirling flow field. Although numerical methods provide insights into three-dimensional mechanisms of swirling flow, their predictions of key features such as recirculation zones and swirling jets are often unsatisfactory due to inherent anisotropy and the isotropic nature of the Boussinesq hypothesis. Experimental methods, such as hot wire, laser Doppler velocimetry, and planar particle image velocimetry (PIV), offer more accurate reference data but are limited by sparse or planar observations. In this study, considering the outperformed capability of solving inverse problems, physics-informed neural network (PINN) was adopted to reconstruct the mean swirling flow field based on limited experimental observations from two-dimensional and two-component (2D2C) results. It was found that adding partial information characterizing the swirling flow, such as the swirling jet, could significantly improve the reconstruction of flow field. In addition, film cooling effectiveness was the key variable to evaluate the film cooling performance, which was relatively measurable in the scalar field. To further improve the accuracy of the reconstruction, the multi-source strategy was adopted into the neural network, where the film cooling effectiveness (FCE) of the effusion plate was imported as the scalar source. It was found that the prediction of the flow field near the target plate was improved, where the highest error reduction could reach 76.5%. Finally, through the reconstructed three-dimensional vortex distribution, it was found that swirling flow vortex structures near the swirler exit had a significant impact on cooling effectiveness, causing a non-uniform cooling distribution. This study aims to diagnose the three-dimensional swirling flow field with deep learning by leveraging limited experimental data and deepen the understanding of effusion cooling under swirling flow condition so that obtains a more accurate reference in the design of thermal protection modules.

Experimental Observations on Film Cooling Effectiveness on Aerodynamic Heating in Hypersonic Flow

An experimental campaign was carried out to examine the effectiveness of film cooling on aerodynamic heating at hypersonic velocities. A flat-faced cylindrical model was tested at flow velocity of 2.5 km/s in a hypersonic shock tunnel. The test model was equipped with temperature sensors for convective heat flux measurement. The coolant gas, nitrogen, was injected through a 2 mm orifice housed at the center of the test model at different pressure ratios. Two different flow regimes were observed, steady and unsteady, with the steady flowfield having overall reduction in heat transfer compared to the unsteady one.

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.

Effects of vortex generators and pulsation on compound angle film cooling

Compound angle film cooling is widely applied in modern turbine engines to provide reliable protection for the components. As passive and active methods, vortex generators enhance the thermal performance significantly, and pulsation coolant jet increases film cooling effectiveness when the compound angle is low. In the current study, the effects of vortex generators, pulsation, and their interactions have been investigated in film cooling with compound angles of 0 deg, 30 deg, and 60 deg. Cooling efficiency distribution is analyzed considering the coolant amount and separation condition. The downwash anticounter-rotating vortex pair (anti-CRVP) generated by vortex generators effectively increases film cooling effectiveness (FCE) in compound angle film cooling. Pulsation enhances the cooling performance by reducing the coolant separation when the lateral angle to the freestream is 0 deg; however, the performance is poor with 30 deg and 60 deg. Vortex generators are less efficient in enlarging the film coverage when pulsed jets are employed due to less coolant close to the surface. Moreover, mean and instantaneous vortex structures are illustrated to show the evolvements and interactions of the anti-CRVP, the counter-rotating vortex pair, the separated wakes, and the asymmetric vortex.

Film-Cooling effectiveness of combined hole slot geometries on a turbine rotor blade surface

A film cooling experiment was conducted for a novel hole structure, namely, the combined hole slot. The basic geometric shape of the combined hole slot is connecting adjacent sidewalls of two parallel diffuser holes internally to form a continuous slot exit. Two combined hole slot geometries constructed based on circular and vertical slot cross-sections were tested and compared with a conventional fan-shaped hole. The film-cooling effectiveness was measured on a turbine rotor blade surface using a linear cascade and pressure sensitive paint technique. The mainstream Reynolds number was 520,000 based on the cascade outlet velocity and the axial chord length. The mainstream turbulence intensity was 3.6%. Coolant-to-mainstream density ratios of 1.5 and 1.0 were tested under five blowing ratios ranging from 0.5 to 2.5. Although an obvious bifurcation occurs at the hole exit, both combined hole slot geometries remarkably increase the film-cooling effectiveness relative to that of the fan-shaped hole, in which the case based on a vertical slot cross-section performs better than the case based on a circular cross-section, regardless of the density ratio. In comparison, the increase in film effectiveness on the suction surface due to the combined hole slot geometries is greater than that on the pressure surface, especially for medium to high blowing ratios. The combined hole slot geometry is generally more suitable for using branch holes with a weak lateral expansion, in which a tighter merging effect can be achieved at the centerline.

Film Cooling Effectiveness and Flow Structures of Butterfly-Shaped Film Cooling Hole Configuration

Abstract Film cooling is applied to various high-temperature components of a gas turbine and greatly affected by the shape of the injection hole. Numerous studies have been conducted to investigate various film cooling structures to increase the film cooling effectiveness by inducing coolants toward the surface. In this study, a butterfly-shaped film cooling hole configuration is proposed, which consists of a pair of asymmetric fan-shaped holes installed adjacently. This novel structure enhances film cooling effectiveness by utilizing the interaction between coolant jets from the asymmetrical film cooling holes. In this study, the film cooling effectiveness of the butterfly-shaped hole configuration was experimentally characterized under low-speed conditions on a flat plate, and was compared with that of widely used fan-shaped holes. Also, a numerical study was conducted to understand and visualize the flow structures generated by the butterfly-shaped film cooling configuration. As a result, an anti-vortex structure was observed near the hole exit. The film cooling effectiveness of the butterfly-shaped hole configuration was also compared with that of the fan-shaped hole. The film cooling effectiveness of the butterfly-shaped hole configuration was 1.9% higher than that of the optimized fan-shaped film cooling hole at its optimized condition. It was also found that the film cooling effectiveness increased with the blowing ratio for the butterfly-shaped hole configuration. It is expected that the film cooling effectiveness can be further improved by adjusting the shape parameters of the butterfly-shaped hole.

ENHANCED FILM COOLING EFFECTIVENESS USING TURBULENCE PROMOTER IN COOLING HOLE

The film cooling is investigated experimentally and numerically for gas turbine blade cooling. Film cooling performance is vital for safe operation of the blades. By calculating the film cooling effectiveness, the effect of wall roughness through a turbulence promoter inside the cooling hole is evaluated. A turbulence promoter of aspect ratio &lt;i&gt;K&lt;/i&gt; &amp;#61; &lt;i&gt;L/h&lt;/i&gt; from 6 to 14, varying pitch &lt;i&gt;L&lt;/i&gt; with constant height &lt;i&gt;h&lt;/i&gt;, is introduced. Blowing ratio range from &lt;i&gt;M&lt;/i&gt; &amp;#61; 0.5 to 2 and fixed mainstream Reynolds number, Re &amp;#61; 2.0 &amp;times; 10&lt;sup&gt;5&lt;/sup&gt; is tested. Smooth cooling hole and flat plate surface with angle of 35&amp;deg; is set up in a wind tunnel to compare the results and validate the CFD approach assessing the turbulence models &lt;i&gt;k-&amp;epsilon;&lt;/i&gt; renormalized group theory (RNG) and Reynolds stress model (RSM). A swirl flow develops inside the cooling hole due to installing the turbulence promoter. The coolant discharge is modified in turbulence parameters improving momentum and heat transfer rates compared to the normal case. This has an impact on the film cooling performance. The results indicate film cooling effectiveness and surface protection enhancement, obtained for the best aspect ratio and blowing ratio. Coolant wall roughness may be used to improve the design of coolant hole and to reduce the number of orifices needed for safe gas turbine blades.

Experimental Study on the Improvement of the Film Cooling Effectiveness of Various Modified Configurations Based on a Fan-Shaped Film Cooling Hole on a Flat Plate

Experimental Study on the Improvement of the Film Cooling Effectiveness of Various Modified Configurations Based on a Fan-Shaped Film Cooling Hole on a Flat Plate

Experimental Study on the Improvement of Film Cooling Effectiveness of Various Modified Configurations Based on a Fan-Shaped Film Cooling Hole on an Endwall

Several studies have previously been conducted to improve the cooling performance of film cooling. However, most of the research has conducted experiments with film cooling holes on flat plates, and thus, the results of these studies do not encompass the influence of the complex mainstream behavior within the turbine passage on film cooling. In this study, three different film cooling hole configurations were installed on the endwall of a turbine linear cascade to measure adiabatic film cooling effectiveness and evaluate cooling performance. The film cooling holes compared in the experiment for film cooling effectiveness were a 7-7-7 fan-shaped hole (Baseline), a Baseline with a double-step structure at the hole exit (Staircase), and a Baseline with an additional expanded passage at the hole leading edge (Compound Expansion). A total of nine holes were manufactured on the turbine endwall to assess film cooling performance, as various factors, such as mainstream acceleration, secondary flow within the turbine passage, and so on, can influence film cooling. Adiabatic film cooling effectiveness was measured using the pressure-sensitive paint (PSP) technique. Mass flow ratios ranging from 0.25% to 1.25% of the mass flow rate of a single turbine passage were supplied to the plenum chamber within the test rig. As a result, all experimental results confirmed the impact of secondary flow within the turbine passage on film cooling. In the case of the Staircase, it exhibits an overall cooling trend similar to the Baseline. It shows small cooling performance degradation compared with Baseline due to lift-off, and its double-step structure laterally expanding results in better cooling performance at high mass flow ratio (MFR) conditions. For the Compound Expansion, at low MFR, the momentum of the coolant is lower compared with other configurations, leading to lower cooling performance due to the influence of secondary flow. However, at high MFR, the Compound Expansion provides wider protection compared with other hole geometries and shows high cooling performance.

Study on Additive Effect of Film Cooling Effectiveness in Two Rows of Fan-Shaped Holes

Abstract The coolant jet interaction has a great influence on the superposition prediction of multirow film cooling. Although there have been many efforts to reveal the mechanics of additive effect in multirow film cooling, the available knowledge about developing the superposition method is still limited. The present work examines the film cooling effectiveness in two rows of fan-shaped holes by pressure sensitive paint technique, at the blowing ratios of 0.5–2.0 and the density ratio of 1.0. It is found that the impact of upstream flow on the downstream cooling film is reflected in the variation of turbulence intensity. The enhanced turbulence intensity is detrimental to the downstream film cooling effectiveness especially at the far away region. The mixing of upstream flow and coolant ejection starts at the leading edge of the hole exit. Thus, the streamwise width of the hole exit should be taken into consideration for better predicting the film cooling effectiveness around the holes. The cause of additive effect is that the coolant ejection at the second row affects the local mainstream entrainment. Then, a new correction factor, which characterizes the influence of coolant ejection on the mainstream entrainment of the upper row, is proposed for improving the classical Sellers method. The final result shows a good agreement with experimental data.

Numerical analysis of flow and heat transfer characteristics of unsteady film cooling of HPT shroud under high-speed rotation sweeping of the rotor blade

3D URANS simulations were performed to investigate the influence of high-speed blade rotation sweeping effect on the unsteady multi-hole film cooling performance of a high-pressure turbine (HPT) shroud. The blade rotation speed was 10,000 rpm and 15,000 rpm, the associated Strouhal numbers (St) were 2.35 and 3.52, and the blow ratio (M) was varied from 0.5 to 1.5. The instantaneous and time-averaged characteristics of film outflow mass rate, pressure distribution, film cooling effectiveness (FCE), and convective heat transfer coefficient (CHTC) of the HPT shroud with and without impinging jets were determined. The results show that the St and M have a substantial impact on the distribution of the FCE and CHTC on the shroud surface. At a high St and M, the FCE on the shroud surface is greater and more sensitive to the rise in M. As St increases, the suppression effect of the tip leaking vortex (TLV) on film and FCE on the shroud surface improved. The increased impingement jet effect has minimal impact on the FCE of the shroud. Under a high-speed blade rotation sweep, the net heat flux on the shroud wall with impingement/film cooling falls by approximately 22% when compared to the uncooled shroud surface.

Experimental and numerical investigation on the metering and diffuser length effects of a laidback fan-shaped film cooling hole

In this study, the metering and diffuser length effects of a laidback fan-shaped (LBF) hole on the film cooling effectiveness (FCE) were investigated. The FCE was measured using the pressure sensitive paint (PSP) technique, with a density ratio (DR) of 2.0 and five blowing ratios (BR). The experimental results indicated that the FCE increased as the diffuser length increased. However, when the metering length ratio increased beyond the critical length, the coolant was shifted to a specific side of the diffusion section, resulting in a lower FCE. A computational analysis was performed to investigate the coolant shift at longer metering lengths. Numerical results indicate that the center velocity of the cylinder section is higher for the longer metering length cases, which leads to a large separation bubble in the diffusion section, these resulted in the increase of the flow unsteadiness and instability at the hole exit.