Cooling Holes Research Articles

Enhancement of Aeroengine Combustion Chamber Air Cooling Holes Design for Emission Reduction and Pattern Factor Control

The results of a comprehensive numerical analysis of temperature uniformity and NOx and CO emission predictions in an aeroengine annular combustor liner, conducted by geometrical modifications through design changes in primary cooling air, including effusion cooling holes, are shown in the paper. A total of five geometric configurations were studied using computational fluid dynamics (CFD) simulations in ANSYS CFX. The adopted combustion model was sort of a combination of the finite-rate (i.e., chemistry, chemical processes) and eddy dissipation model (FRC/EDM). Further, the combustion of liquid kerosene (C10H22) with air, subsequent to the evaporation of fuel droplets, was simulated, and the Rosin-Rammler droplet size distribution was used in spray modeling for accurate depiction of fuel atomization. Both thermal and prompt formation mechanisms of NOx were considered, while k-ε model for turbulence was adapted to capture the nature of emission. An annular combustion chamber of realistic dimensions with a double radial air swirler was modelled in 3D CAD to undertake this study for good, tangible results. Built contour plots allowed to analyze the temperature distribution and NOx concentration along the axis from the center of the injector. Charts on the pattern factor, temperature, and NOx and CO concentrations at the outlet of the combustor served as performance metrics. The simulation was implemented with a two-step chemical kinetics scheme for kerosene combustion with the P1 radiation model, which would give an accurate thermal radiation prediction. One of the major objectives of this research is to compare the CFD results at the combustor outlet with gas dynamic and thermodynamic calculations that have been carried out using AxStream software at the Department of Aeroengine Design, Kharkiv Aviation Institute. It is important to emphasize that the mean deviation of gas dynamic results obtained from AxStream and CFD simulation results was found to be insignificant, hence the CFD approach has been validated. The results testify that redesigning the combustor liner, especially in the design related to primary and effusion cooling holes, drastically reduced NOx and CO emissions. Also, these design modifications have helped in reducing or improving temperature uniformity at the combustor outlet, either way enhancing combustion efficiency and performance.

In-Hole Measurements of Flow Inside Fan-Shaped Film Cooling Holes and Downstream Effects

The study of low-speed jets into crossflow is critical to the performance of gas turbines. Film cooling is a method to maintain manageable blade temperatures in turbine sections while increasing turbine inlet temperatures and turbine efficiencies. Initially, cooling holes were cylindrical. Film cooling jets from these discrete round holes were found to be very susceptible to jet liftoff, which reduces surface effectiveness. Shaped holes have become prominent for improved coolant coverage. Fan-shaped holes are the most common design and have shown good improvement over round holes. However, fan-shaped holes introduce additional parameters to the already complex task of modeling cooling effectiveness. Studies of these flows range in hole lengths from those found in actual turbine blades to very long holes with fully developed flow. The flow within the holes themselves is difficult to study as there is limited optical access. However, the flow within the holes has a strong effect on the resulting properties of the jet. This study presents velocity and vorticity fields measured using high-resolution magnetic resonance velocimetry (MRV) to study three different fan-shaped hole geometries at two blowing ratios. Because MRV does not require line of sight, it provides otherwise hard-to-obtain experimental data of the flow within the film cooling hole in addition to the mainflow measurements. By allowing measurement within the cooling hole, MRV shows how a poor choice of diffuser start point and angle can be detrimental to film cooling if overall hole length and cooling flow velocity are not properly accounted for in the design. The downstream effect of these choices on the jet height and counter-rotating vortex pair is also observed.

High-Cycle Fatigue Fracture Behavior and Stress Prediction of Ni-Based Single-Crystal Superalloy with Film Cooling Hole Drilled Using Femtosecond Laser

A high-temperature, high-cycle fatigue test was conducted on a nickel-based single-crystal superalloy with a pore structure. Optical and scanning electron microscopy were utilized to examine the crack propagation paths and fatigue fracture surfaces at the macro and micro scales. The analysis of crack initiation and propagation related to the pore structure facilitated the development of a crack shape factor reflecting these distinct fracture behaviors. Predictions about the high-cycle fatigue stress experienced by the specimen were made, accompanied by an error analysis, providing critical insights for precise stress calculations and structural optimization in engine blade design. The results reveal that high-cycle fatigue cracks originate from corner cracks at pore edges, with the initial propagation displaying smooth crystallographic plane features. Subsequent stages show clear fatigue arc patterns in the propagation zones. The fracture surface exhibits the significant layering of oxide layers, primarily composed of NiO, with traces of CoO displaying columnar growth. AL2O3 is predominantly found at the interfaces between the matrix and oxide layers. Short and straight dislocations near the oxide layers and within the matrix suggest that dislocation multiplication and planar slip dominate the slip mechanisms in this alloy. The orientation of the fracture surface is mainly perpendicular to the load direction, with minor inclined facets in localized areas. Correlations were established between the plastic zone dimensions at the crack tips and the corresponding fatigue stresses. Without grain boundaries in single-crystal alloys, these dimensions are easily derived as parameters for fatigue stress analysis. The selected crack shape factor, “elliptical corner crack at pore edges”, captures the initiation and propagation traits relevant to porous structures. Subsequent calculations, accounting for the impact of oxide layers on stress assessments, indicated an error ratio ranging from 1.00 to 1.21 compared to nominal stress values.

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

Consideration of Cooling Hole Configuration Taking into Account Aerodynamic Losses due to Film Cooling on the Endwall of Gas Turbine Blades

Consideration of Cooling Hole Configuration Taking into Account Aerodynamic Losses due to Film Cooling on the Endwall of Gas Turbine Blades

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.