Cooling Holes Research Articles

Numerical Stress Analysis of TBCs at Film Cooling Hole's Edge based on TBC-Film Cooling System

Numerical Stress Analysis of TBCs at Film Cooling Hole's Edge based on TBC-Film Cooling System

Geometric and Flow Characterization of Additively Manufactured Turbine Blades With Drilled Film Cooling Holes

Abstract As designers investigate new cooling technologies to advance future gas turbine engines, manufacturing methods that are fast and accurate are needed. Additive manufacturing facilitates the rapid prototyping of parts at a cost lower than conventional casting but is challenged in accurately reproducing small features such as turbulators, pin fins, and film cooling holes. This study explores the potential application of additive manufacturing and advanced hole drill methods as tools to investigate cooling technologies for future turbine blade designs. Data from computed tomography scans are used to nondestructively evaluate each of the cooling features in the blade. The resulting flow performance of these parts is further related to the manufacturing through benchtop flow testing. Results show that while total blade flow is consistent for all additively manufactured cooled blades, flow through smaller regions of the blades shows variations. Shaped film cooling holes manufactured using a high-speed electrical discharge machining method are within tolerance in the metering section but do not expand at the specified angle in the diffuser even though design tolerances are met. In contrast to high-speed EDM, conventional EDM holes are undersized throughout the length of the hole. Due to the additive manufacturing process, the surface roughness was higher on the additively manufactured parts in the current study than has been previously reported for surface roughness of commonly used cast components. The roughness results show high levels on thin walls, particularly at the trailing edge as well as on downskin surfaces. Internal surface roughness is higher than external roughness at most locations on the blade. The results of this study confirm that additive manufacturing along with advanced hole drilling techniques offers faster development of blade cooling designs.

A Review of the Machining of the Film Cooling Holes with Thermal Barrier Coatings Through Non-traditional Machining Methods

A Review of the Machining of the Film Cooling Holes with Thermal Barrier Coatings Through Non-traditional Machining Methods

A SUPERPOSITION TECHNIQUE FOR PREDICTING OVERALL EFFECTIVENESS WITH MULTIPLE SOURCES OF INDIVIDUALLY CHARACTERIZED INTERNAL AND EXTERNAL COOLANT FLOWS

Abstract Gas turbine engines rely on intricate cooling schemes to protect turbine components from the high temperature freestream gas. Determining the optimal placement of film cooling holes while remaining mindful of the consumption of coolant air is therefore a design challenge. For several decades a method of superposition has been in use that provides a reasonable prediction of the adiabatic effectiveness in a region influenced by multiple rows of film cooling holes acting together if the adiabatic effectiveness distributions resulting from individual rows are already known. This classical superposition technique has been used to provide a first cut at determining where coolant holes might be placed. A shortcoming of the method is that it strictly applies only to prediction of adiabatic effectiveness which is indicative of the adiabatic wall temperature, not the actual surface temperature. The overall effectiveness, which is indicative of the actual surface temperature is somewhat more complicated as it is influenced not only by the external cooling, but also internal cooling. In the present work, a method of superposition of overall effectiveness data is proposed, allowing prediction of the actual surface temperature distribution on a component. Experimental results show that it is possible to use knowledge of the overall effectiveness distributions resulting from individual internal cooling and associated film cooling holes acting alone to determine how they are likely going to act when combined. This technique shows promise for a turbine designer to use superposition for actual surface temperature prediction and therefore higher quality initial cooling designs.

Numerical Analysis of Film Cooling Flow Dynamics and Thermodynamics for Perfect and Imperfect Cooling Holes

Numerical Analysis of Film Cooling Flow Dynamics and Thermodynamics for Perfect and Imperfect Cooling Holes

Overall Cooling Effectiveness With Internal Serpentine Channels and Optimized Film Cooling Holes

Abstract The overall cooling effectiveness for gas turbine airfoils is a function of the combined cooling due to internal cooling configurations and film cooling configurations. Typically, film cooling configurations are evaluated independent of the cooling effects of the internal feed channels, generally based on adiabatic effectiveness measurements. In this study, we consider the coupled effects of internal cooling and film cooling configurations through measurements of overall cooling effectiveness for film cooling holes fed by a coflow/counterflow channel and a serpentine channel. A film cooling hole designed by adjoint optimization techniques (X-AOpt) is compared to a standard-shaped hole with 7 deg forward and lateral expansions (7-7-7 SI). Experiments without film cooling showed that the serpentine channel had 35–50% greater overall cooling effectiveness than the straight, coflow channel. Experiments with the X-AOpt hole combined with a serpentine channel showed an area-averaged overall cooling effectiveness of ϕ¯¯=0.58, which was a 70% increase compared to the overall cooling effectiveness of the serpentine channel without film cooling. When the X-AOpt hole was fed with a coflow channel with similar coolant mass flowrate, the overall cooling effectiveness was ϕ¯¯=0.44, i.e., 30% lower than when using the serpentine channel. Interestingly, adiabatic effectiveness measurements with the X-AOpt holes showed a more uniform hole-to-hole performance when using the serpentine channel compared to the coflow channel.

Advances in Aeroengine Cooling Hole Measurement: A Comprehensive Review.

Film cooling technology is of great significance to enhance the performance of aero-engines and extend service life. With the increasing requirements for film cooling efficiency, researchers and engineers have carried out a lot of work on the precision and digital measurement of cooling holes. Based on the above, this paper outlines the importance and principles of film cooling technology and reviews the evolution of cooling holes. Also, this paper details the traditional measurement methods of the cooling hole used in current engineering scenarios with their limitations and categorizes digital measurement methods into five main types, including probing measurement technology, optical measurement technology, infrared imaging technology, computer tomography (CT) scanning technology, and composite measurement technology. The five types of methods and integrated automated measurement platforms are also analyzed. Finally, through a generalize and analysis of cooling hole measurement methods, this paper points out technical challenges and future trends, providing a reference and guidance for forward researches.

Aerothermal Optimization of Film Cooling Hole Locations on the Squealer Tip of an HP Turbine Blade

Abstract This article presents a time-efficient method to optimize the positions of film cooling holes on a gas turbine blade's squealer tip for cooling and aerodynamic performance. A computational approach is employed for the optimization, including validations against experiments. Five discrete film cooling holes are considered, and two different blowing ratios of 0.4 and 1.0 are studied. The positions of cooling holes on the tip along the tangential direction are varied as the input parameters of optimization. The multi-objective optimization uses an algorithm with an artificial neural network for fast fitness function predictions. The best cooling configuration found by the optimization achieves a 13.43% reduction in total heat flux and a 0.4% increase in aerodynamic loss when the blowing rate is 1.0. Including the casing relative motion in the computations results in a total pressure loss coefficient increase of about 8% for both blowing ratios. For M = 1.0, imposing the casing's motion results in a 10.2% reduction in total heat transfer to the tip compared to the stationary casing. For the lower blowing rate of 0.4, the total heat flux reduction to the tip is 12.0% because of the imposed casing motion. Hence, the cooling effectiveness can be improved by employing the particular position optimization method presented in this study. The results suggest that experimental and computational heat transfer studies on cooled turbine blade tips, especially in cascade arrangements, need to consider the relative motion of the blade tip.

Data-driven framework for prediction and optimization of gas turbine blade film cooling

Film cooling is a crucial technique for protecting critical components of gas turbines from excessive temperatures. Multiparameter film cooling optimization is still relatively time-consuming owing to the substantial computational demands of computational fluid dynamics (CFD) methods. To reduce the computational cost, the present study develops a data-driven framework for predicting and optimizing the film cooling effectiveness of high-pressure turbines based on deep learning. Multiple rows of cooling holes located on the pressure surface of the turbine blade are optimized, with the coolant hole diameter, the incline angle, and the compound angle as design parameters. A conditional generative adversarial network model combining a gated recurrent unit and a convolutional neural network is designed to establish the complex nonlinear regression between the design parameters and the film cooling effectiveness. The surrogate model is trained and tested using independent CFD results. A sparrow search algorithm and the well-trained surrogate model are combined to acquire the optimal film cooling parameters. The proposed framework is found to improve multi-row film cooling effectiveness by 21.2% at an acceptable computational cost.

Achieving Best Turbine Blade Cooling Efficiency from a Cylindrical Cooling Hole with Reverse-Vortex Generator Using Numerical Simulation

This investigation bargains mathematical investigation of the impact of kidney vortex generation (KVG) on the film cooling execution of a round cooling opening which has measurement (d = 20 mm) on a level plate. The collection between the stream cooling air and standard hot gases will result kidney vortex which will wipe out the film cooling adequacy. Two kind of reverse vortex generator (RVG) with various statures (H = 0.5 d and 0.25 d) are intended to eliminate the kidney vortex and examine best film cooling adequacy, where every one of them is mounted to the level plate upstream of the cooling opening by changing its parallel positions (A = 0.0 d, 0.25 d, 0.5 d and 0.75 d) concerning the opening community line and for each kind has diverse distance regard to opening focus line. The changing of blowing proportion (BP = 0.5-1.0) was considered in this investigation. The simulation model was used to simulate the formation of film cooling using a shear stress transfer (SST) model by Workbench 15 software. The outcomes have been introduced in terms of laterally averaged and maximum value of film cooling comparison diagrams, vorticy on x/d = 3.0 and film cooling distribution which clarified how the outcomes acquired. The Cases 04 and 07 gave the best situations for RVG where case 04 gave wide covered for laterally average film cooling efficiency (η = 0.35), while case 07 gave the highest expansion distribution and the maximum value for film cooling efficiency (η = 0.66). The best results are obtained when the blowing ratio (BP = 0.5) due to a lower velocity of cooling air which will made best interaction with the hot air (main stream)which will represent the hot gases in real case.

Experimental and Numerical Investigation of Cylindrical and Shaped Cooling Holes With Forward and Reverse Injection

The present work proposes a suitable injection hole configuration of cylinder and laidback fan-shaped with forward and reverse direction for film cooling of a gas turbine blades application, based on experimental and numerical analysis. The experimental study is conducted for cylindrical hole at blowing ratio (1), injection angle (35°), and density ratio (1.2). The numerical study is performed for a wide range of operating parameters such as blowing ratios on (1-3), density ratios (2.42), mainstream flow Reynolds number as 4000 based on the hydraulic diameter of wind tunnel channel and, injection angle (35°) with the effect of forward and reverse injection of laidback fan shaped. The present study reveals that the formation of kidney vortices mitigated for reverse-shaped holes (secondary air is injected such that its axial velocity component is in the reverse direction to that of the mainstream) results in higher cooling performance with respect to forward-shaped holes. The coolant coverage is likewise more consistent and higher in the lateral direction compared to the forward injection.

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.

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 <i>K</i> = <i>L/h</i> from 6 to 14, varying pitch <i>L</i> with constant height <i>h</i>, is introduced. Blowing ratio range from <i>M</i> = 0.5 to 2 and fixed mainstream Reynolds number, Re = 2.0 × 10<sup>5</sup> is tested. Smooth cooling hole and flat plate surface with angle of 35° is set up in a wind tunnel to compare the results and validate the CFD approach assessing the turbulence models <i>k-ε</i> 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.

A novel design for film-cooling: Cooling holes with inlet groove

A novel design for film-cooling: Cooling holes with inlet groove

Adaptive Drilling of Film Cooling Holes of Turbine Vanes Based on Registration of Point Clouds

The manufacture of film cooling holes on turbine guide vanes of an aero engine suffers from the contour errors resulting from a casting process. It is therefore necessary to adjust positions and orientations of cooling holes for each vane. This paper proposes an adaptive compensation machining method based on a combination of a principal component analysis iterative closest point (PCA-ICP) registration method and an orthocenter mapping method. Firstly, a registration is carried out between an ideal point cloud and a measured point cloud by a laser measurement machine to obtain a transformation matrix. Then, the exit center of a cooling hole is obtained by an orthocenter mapping method. Simulation results show that the machining errors of vanes can be improved from <inline-formula xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink"><tex-math notation="LaTeX">$\pm$</tex-math></inline-formula> 0.1 mm to <inline-formula xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink"><tex-math notation="LaTeX">$\pm$</tex-math></inline-formula> 0.03 mm. Machining tests show that with a combination of the PCA-ICP registration and orthocenter mapping methods, the cooling holes are drilled adaptively to meet the machining requirements.

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 comparison of the indexable drilling with different U drills on hole quality

U drills are estimated to account for around 53% of all drilling equipment’s market share. This study compared and contrasted the impacts of the balance angle, the U drill length/diameter ratio, and the extra coolant hole opened in the chip evacuation channel on the hole quality when drilling AA 7075-T651 material. Hole quality was investigated by deviation from diameter, deviation from circularity, deviation from cylindricality, surface roughness, forces, and torque. Three cutting speeds, three feed rates, and four distinct U drills were chosen as input parameters. The studies revealed that when the U drill’s length/diameter ratio increased, the hole quality deteriorated. The best hole quality was nonetheless achieved by the tool with the additional coolant hole. The higher balance angle resulted in improved hole quality. However, it was discovered after looking at the chips and inserts that the chatter only happened while drilling with a U drill that had an additional coolant hole. While the extra coolant hole opened to the chip evacuation channel, it affected the hole quality positively but caused chatter, which negatively affected the inserts. The reasons for these results are discussed in detail.

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.

Comparisons of Overall Performance Among Double-Jet Film Cooling Holes, Cylinder Holes, and Fan-Shaped Holes

Abstract In this paper, the film-cooling effectiveness (η) and heat transfer coefficient (h) of different film hole geometries are investigated, including double-jet film cooling (DJFC) holes, streamwise cylindrical holes, and fan-shaped holes, both experimentally and numerically. Results reveal that when the blowing ratio is less than 1.0, the DJFC holes have the highest η and the highest h, as well as the highest net heat flux reduction (NHFR). However, a higher blowing ratio (&amp;gt;1.0) leads to a quickly decreasing NHFR of DJFC holes. The asymmetric antikidney vortex and the high turbulent kinetic energy (TKE) are dominant in the performance of the DJFC holes. Owing to medium effectiveness and the lowest heat transfer coefficient, the fan-shaped holes possess the highest net heat flux reduction at M = 2.0 although the value is negative. The relatively weak kidney vortex and the low TKE can explain the phenomena. The cylindrical holes have the lowest η and the lowest NHFR due to the kidney vortex and relatively higher TKE.