David G. Bogard

Abstract Compressible flow fields associated with engine realistic high-speed conditions can have a significant impact on the performance of shaped film cooling holes in gas turbines. Such flow fields are dependent upon a variety of parameters, notably the stagnation temperature ratio, an analog to the density ratio, and the mainstream Mach number. In previous high-speed computations using 7-7-7 shaped hole geometry, cases with a higher stagnation temperature ratio performed significantly worse than those with a lower stagnation temperature ratio, at the same blowing ratios. The same computational work, in addition to a preliminary experimental study, demonstrated that 7-7-7 cases with elevated mainstream Mach numbers perform significantly worse than those with low mainstream Mach number. In the present study, experiments were performed with 7-7-7 shaped film cooling holes across a wide range of conditions. The stagnation temperature ratio was varied from 0.6 to 0.8, and the mainstream Mach number was varied from 0.15 to 0.50. The results confirmed that the stagnation temperature ratio has a significant impact on the performance at high speeds, with higher performance occurring at a lower stagnation temperature ratio, or by proxy higher density ratio. Furthermore, cases with an elevated mainstream Mach number performed significantly worse than those with low mainstream Mach number. For all cases, performance was scaled with both blowing ratio and pressure ratio, and the implications of such scaling are discussed.

Abstract The goal of this study was to develop a highly cooled, turbine nozzle designed for use in extremely high temperatures exhausting from a hydrogen-fueled combustor. This design was based on the nozzle being manufactured using metal additive manufacturing techniques. The increasing fidelity of additive manufacturing enables the fabrication of complex geometries, allowing the implementation of more optimized cooling schemes. The present study details the design of a symmetric airfoil, or strut, mimicking that of a turbine vane. In particular, the final design featured novel film cooling holes recently developed in our laboratory, and an array of impinging jets and a lattice structure of pin fins were the mechanisms responsible for internal cooling. Reynolds-averaged Navier–Stokes (RANS) computational simulations were used to determine coolant flow within the airfoil and exhausting film cooling holes for the final strut design. For validation of the computational predictions, an experimental model of the final design was fabricated and tested in a low-speed wind tunnel facility. Results will be shown comparing computational predictions and experimental measurements. In addition, the strut was simulated under engine realistic temperatures, 1750K, and approach velocities equivalent to a mainstream Mach number of Ma∞=0.039. The design was found to successfully keep the wall temperature sufficiently below the melting temperature of Inconel 718. These conditions match those of a hydrogen combustor facility at Southwest Research Institute, which will be used in future work to validate the results of the present study.

Abstract This study evaluated the overall cooling effectiveness of a newly designed turbine airfoil using optimized internal and external cooling features. A symmetric airfoil, which incorporated advanced cooling schemes enabled by additive manufacturing, was designed for testing in a high-temperature hydrogen-fueled combustor facility. The design of this airfoil is described in a companion study. This article describes the experimental and computational evaluation of the new design to ensure that the test model would have sufficient cooling to operate in the high-temperature combustor. Tests were performed in a low-temperature wind tunnel facility, matching the Mach number and Reynolds number of the combustor test facility. Further insight into the cooling performance and analysis of the experimental results was obtained using computational predictions. Experimental testing included a low-conductivity model that was used to estimate adiabatic effectiveness for film cooling configurations. A high conductivity model, which matched the Biot number expected for operation in the combustor facility, was tested to obtain overall cooling effectiveness. Experimental results showed adiabatic effectiveness was generally above η¯=0.3, though there was a small region that fell below η¯. Overall cooling effectiveness was quite uniform ranging from ϕ¯=0.54to0.63 over the complete airfoil. Reynolds-averaged Navier–Stokes (RANS) computations were consistent with the experimental trends. The RANS predictions were particularly useful in showing how conduction effects in the low-temperature model were affecting the measurements of adiabatic effectiveness. The experiments and computational predictions verified that the design intent to have a relatively uniform overall effectiveness of greater than ϕ¯=0.55 was achieved. The study also identified deficiencies in the cooling design that could potentially be addressed leading to improved overall cooling effectiveness.

Abstract Adjoint based shape optimization has the potential to produce real gains in the performance of gas turbine film cooling geometry. This has been demonstrated in previous studies with incompressible mainstream Mach number, using both computational and experimental evidence. Starting from a 7-7-7 shaped film cooling hole as a baseline, the present study investigated such optimization routines using an elevated mainstream Mach number of Ma∞ = 0.75, and is the first of its kind. At this elevated Mach number, supersonic flow and shockwaves were predicted within the 7-7-7 baseline hole. How such behaviors affected the optimization was of particular interest. In addition to the optimization at elevated mainstream Mach number, a control optimization was performed at an incompressible mainstream Mach number of Ma∞ = 0.15. Both the “high-speed” and “low-speed” optimizations were performed using a hole Reynolds number ReD = 10,000, a blowing ratio M = 2.00, and a stagnation temperature ratio TR° = 0.50. Although the resulting optimizations were similar, significant differences were observed and are discussed. After generating each optimization, the resulting geometries were simulated across a wide range of conditions. Computationally, the predicted performance of both geometries was significantly higher than that of the baseline 7-7-7. Each geometry was also additively manufactured and tested experimentally at mainstream Mach numbers of Ma∞ = 0.15 and Ma∞ = 0.50, for a limited range of conditions. Peak experimental performance of both optimizations was greater than the baseline 7-7-7 at Ma∞ = 0.15, but for the range of blowing ratio tested at Ma∞ = 0.50, neither optimization had experimental peak performance greater than the baseline 7-7-7.

Abstract The overall cooling effectiveness of a gas turbine airfoil characterizes the conjugate effect of both film cooling and internal cooling on an airfoil in one parameter. Few studies in open literature have explored the overall cooling effectiveness of airfoils utilizing thermal barrier coatings. In this study it was found that the addition of a 0.84D thick TBC to an internally cooled flat plate yields a 130% improvement in the overall cooling effectiveness. This result along with increased reliability of TBCs suggests that TBCs should be considered when determining optimal integrated cooling configurations for an airfoil. RANS computational simulations were used to determine the optimal cooling configuration for a flat plate with a single row of film cooling holes and TBC. A variety of parameters were iterated upon to determine a film cooling configuration that provided a 14% improvement in overall cooling effectiveness relative to the case without film cooling while using the least amount of coolant. Multi-row simulations were also performed in order to determine an optimal streamwise spacing. The major findings of this study reveal that when in the presence of a TBC, a round hole in a trench is the highest performing film cooling configuration. It was also determined that to reach a 14% improvement, the lateral spacing of the film cooling holes must still be the traditional spacing of P/D = 3. Finally, experiments of a select number of the single row configurations simulated were tested experimentally to compare to the RANS simulations.

Abstract The goal of this study was to develop a highly cooled, turbine nozzle designed for use in extremely high temperatures exhausting from a hydrogen fueled combustor. This design was based on the nozzle being manufactured using metal additive manufacturing techniques. The increasing fidelity of additive manufacturing enables the fabrication of complex geometries, allowing the implementation of more optimized cooling schemes. The present study details the design of a symmetric airfoil, or strut, mimicking that of a turbine vane. In particular, the final design featured novel film cooling holes recently developed in our laboratory, and an array of impinging jets and a lattice structure of pin fins were the mechanisms responsible for internal cooling. RANS computational simulations were used to determine coolant flow within the airfoil and exhausting film cooling holes for the final strut design. For validation of the computational predictions, an experimental model of the final design was fabricated and tested in a low-speed wind tunnel facility. Results will be shown comparing computational predictions and experimental measurements. In addition, the strut was simulated under engine realistic temperatures, 1750K, and approach velocities equivalent to a mainstream Mach number of Ma∞ = 0.039. The design was found to successfully keep the wall temperature sufficiently below the melting temperature of Inconel 718. These conditions match those of a hydrogen combustor facility at Southwest Research Institute, which will be used in future work to validate the results of the present study.

Abstract This study evaluated the overall cooling effectiveness of a newly designed turbine airfoil using optimized internal and external cooling features. A symmetric airfoil, which incorporated advanced cooling schemes enabled by additive manufacturing, was designed for testing in a high temperature hydrogen fueled combustor facility. The design of this airfoil is described in a companion study. This paper describes the experimental and computational evaluation of the new design to ensure that the test model would have sufficient cooling to operate in the high temperature combustor. Tests were performed in a low temperature wind tunnel facility, matching the Mach number and Reynolds number of the combustor test facility. Further insight into the cooling performance and analysis of the experimental results was obtained using computational predictions. Experimental testing included a low conductivity model that was used to estimate adiabatic effectiveness for film cooling configurations. A high conductivity model, which matched the Biot number expected for operation in the combustor facility, was tested to obtain overall cooling effectiveness. Experimental results showed adiabatic effectiveness was generally above η¯ = 0.3, though there was a small region that fell below η¯. Overall cooling effectiveness was quite uniform ranging from ϕ¯ = 0.54 to 0.63 over the complete airfoil. Reynolds-Averaged Navier-Stokes (RANS) computations were consistent with the experimental trends. The RANS predictions were particularly useful in showing how conduction effects in the low temperature model were affecting the measurements of adiabatic effectiveness. The experiments and computational predictions verified that the design intent to have a relatively uniform overall effectiveness of greater than ϕ¯ = 0.55 was achieved. The study also identified deficiencies in the cooling design that could potentially be addressed leading to improved overall cooling effectiveness.

Abstract Compressible flow fields associated with engine realistic high-speed conditions can have a significant impact on the performance of shaped film cooling holes in gas turbines. Such flow fields are dependent upon a variety of parameters, notably the stagnation temperature ratio, an analog to density ratio, and the mainstream Mach number. In previous high-speed computations using 7-7-7 shaped hole geometry, cases with higher stagnation temperature ratio performed significantly worse than those with lower stagnation temperature ratio, at the same blowing ratios. The same computational work, in addition to a preliminary experimental study, demonstrated that 7-7-7 cases with elevated mainstream Mach number perform significantly worse than those with low mainstream Mach number. In the present study, experiments were performed with 7-7-7 shaped film cooling holes across a wide range of conditions. Stagnation temperature ratio was varied from 0.6–0.8, and mainstream Mach number was varied from 0.15–0.50. The results confirmed that stagnation temperature ratio has a significant impact on performance at high speeds, with higher performance occurring at lower stagnation temperature ratio, or by proxy higher density ratio. Furthermore, cases with elevated mainstream Mach number performed significantly worse than those with low mainstream Mach number. For all cases, performance was scaled with both blowing ratio and pressure ratio, and the implications of such scaling are discussed.

Abstract In recent studies, a turbine film cooling shaped hole designed by adjoint optimization techniques (X-AOpt) was found to substantially increase film cooling performance with 90% greater adiabatic effectiveness than a reference 7-7-7 shaped hole. Two aspects of the geometry contribute to the high adiabatic effectiveness levels achieved by the X-AOpt holes: the shape of the internal geometry which improves the diffuser performance, and the external protrusions that generate counter-rotating vortices that push the core of the coolant jets towards the wall and spread the coolant laterally. In this study, insight into the relative importance of these two factors was obtained by testing the performance of a row of X-AOpt holes with the protrusions removed from the external surface. Experiments were also performed with varying pitch between X-AOpt holes to determine how the hole spacing affects film cooling effectiveness. Results from the tests of X-AOpt holes without external protrusions showed a 9% increase in adiabatic effectiveness compared to the baseline 15-15-1 holes, while the X-AOpt holes had a 45% higher adiabatic effectiveness than this standard. Experiments were also conducted using X-AOpt holes with varying pitch between holes of P/D = 4.0, 5.45, and 7.0. Results from these experiments showed that a decrease in pitch to P/D = 4.0 provided only a slightly increased adiabatic effectiveness compared to standard P/D = 5.45. However, increasing the pitch to P/D = 7.0 caused a 33% decrease in adiabatic effectiveness compared to the standard spacing. Superposition was found to significantly under predict adiabatic effectiveness for P/D = 4.0 and 7.0, indicating there is a dependence on lateral spacing for the X-AOpt geometry due to vortex interaction.

Abstract Prediction of film cooling performance on turbine airfoils with multiple rows of film cooling holes based on the measured performance of a single row of holes relies on use of superposition analysis. There have been many past studies of superposition predictions of film cooling adiabatic effectiveness with similar hole shapes, either cylindrical or shaped holes. Recently there has also been a study proposing using a superposition technique to predict overall cooling effectiveness for multiple rows of film cooling holes. In this study, superposition predictions for adiabatic effectiveness and overall cooling effectiveness were evaluated for multiple rows of film cooling holes with varying hole geometries. The test configurations used mimicked the influence of upstream showerhead film cooling holes followed by either baseline 7-7-7 holes or an additively manufactured optimized film cooling hole which included external vortex generators that improved attachment of the coolant jets. Results showed the commonly used Sellers superposition model provided reasonable predictions of adiabatic effectiveness for the more standard 7-7-7 holes, but poor predictions for the optimized holes using vortex generators. Accuracy of the predictions of overall cooling effectiveness based on measured adiabatic effectiveness were found to be good.

Abstract Recent evidence suggests that film cooling flows with engine-realistic mainstream Mach number have declined performance in comparison to those with conventional low-speed laboratory conditions. Consideration and understanding of these effects are fundamental to improving film cooling research. This computational study investigates the film cooling performance of a 7-7-7 shaped film cooling hole with respect to varying mainstream Mach number, with constant Reynolds number. The cases studied include mainstream Mach numbers from 0.15 to 0.75, with a fixed, engine realistic, hole Reynolds number of Red=10,100. Significant results are then evaluated against varying stagnation temperature ratio and blowing ratio. The results showed that at a blowing ratio of 1.75, the adiabatic effectiveness declines significantly with increasing mainstream Mach number. The decreased performance is due to supersonic flows and shocks in the film cooling hole that disrupt flow in the diffuser section of the hole. These characteristics are observed across all stagnation temperature ratios considered. In addition to these insights, the study discusses the importance of proper thermal scaling and definition of adiabatic effectiveness when operating at high mainstream Mach number. It is demonstrated that the effects of high-speed flow challenge the efficacy of the conventional parameters used to characterize film cooling performance.

This research program involved the development of improved gas turbine cooling by optimizing the combined internal and film cooling configurations used for turbine blades and vanes. Improved cooling of turbine airfoils allows the gas turbines to be operated at higher temperatures which improves efficiency for gas turbine-based power generation. This also facilitates operation with hydrogen-based fuels which inherently generate higher operating temperatures. Designs of the new cooling configurations were based on the added geometric flexibility that is available when using metal additive manufacturing techniques. When making these designs, adjustments were made to account for imperfections occurring in the metal additive manufacturing process. New designs for film cooling holes were developed using computational adjoint gradient optimization techniques which generated a unique film cooling hole with 70% greater film cooling effectiveness compared to conventional shaped film cooling holes. A major factor in improving the film cooling effectiveness were small external protrusions on either side of the new hole which generated vortices that pushed coolant towards the surface and increased the lateral spreading of the coolant. Our studies also included optimization of film cooling holes that can be built with conventional manufacturing techniques and optimizing the internal coolant channels that supply coolant to the film cooling holes. Various internal coolant channel shapes were investigated and designs that maximized the cooling of the airfoil while minimizing pressure losses in the channel flows were identified. Testing and verification of the performance of the final configurations was done with a transonic wind tunnel facility which allowed testing at engine realistic high velocities. Final testing of the combined internal and film cooling configurations was done by incorporating the configurations in Penn State’s National Experimental Turbine (NExT) vane. These tests confirmed the enhanced overall cooling effectiveness provided by the new cooling configuration designs.

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.

Abstract Cooling components in the hot section of a gas turbine are essential to component durability. Common methods of cooling include rib turbulators in internal passages and film cooling on external surfaces. The holes that produce the film cooling are fed from the internal channels often containing ribs. Consequently, there is an interdependence of internal heat transfer and external film cooling. The purpose of this study was to obtain a better understanding of the interaction of ribs and film cooling. To quantify the cooling performance, the surface temperatures were measured, from which overall effectiveness was calculated. For the experiments, additively manufactured test coupons were made of Inconel 718 to match engine Biot numbers. These test coupons had internal feed channels with and without ribs and had both cylindrical holes and meter-diffuser-shaped holes with 15 deg lateral expansion angles and a 1 deg forward expansion angle. A single rectangular channel was one type of feed channel. The other type of feed channel was individual circular channels, with each circular channel supplying an individual film-cooling hole. The experimental results showed that the circular individual channels have 80% higher baseline overall effectiveness than the single rectangular channels without any film cooling. Ribbed turbulators without film cooling also increased the overall effectiveness by 21% for single rectangular channels and by 29% for the circular individual channels compared to the respective non-ribbed channels. While the film cooling increased the overall effectiveness of all geometries, the single rectangular channels had increased overall effectiveness levels by up to twice that of the no film-cooling case. On average, the single rectangular channels had an 80% improvement from film cooling, whereas the individual channel feeds, on average, had only a 50% improvement, given their high baseline effectiveness levels.

Abstract While modern gas turbine engines operate at hot gas path velocities approaching the speed of sound, few facilities have studied the effects that the flow’s compressibility can have on the adiabatic effectiveness. A new facility at the University of Texas at Austin has been developed to investigate these high Mach number effects and how to appropriately scale laboratory film cooling experiments to engine conditions. This study investigates two film cooling hole geometries, a baseline 7-7-7 shaped film cooling hole and a recent design which has been numerically optimized for increased effectiveness. Both holes are tested at mainstream Mach numbers of 0.25 and 0.50 in a flat plate test section. The optimized hole outperforms the effectiveness of the baseline geometry at all blowing ratios tested, matching the trend in the results of previous studies on these geometries. However, there is a marked decrease in film cooling hole performance as the Mach number is increased.

Abstract Recent evidence suggests that film cooling flows with engine realistic mainstream Mach number have declined performance in comparison to those with conventional low-speed laboratory conditions. Consideration and understanding of these effects are fundamental to improving film cooling research. The proposed computational study investigates the film cooling performance of a 7-7-7 shaped film cooling hole with respect to varying mainstream Mach number, with constant Reynolds number. The cases studied include mainstream Mach numbers from 0.15–0.75, with a fixed, engine realistic, hole Reynolds number of Red = 10, 100. Significant results are then evaluated against varying stagnation temperature ratio and blowing ratio. The results showed that at a blowing ratio of 1.75, the adiabatic effectiveness declines significantly with high mainstream Mach number. The decreased performance is due to supersonic flows and shocks in the film cooling hole that disrupt flow in the diffuser section of the hole. These characteristics are observed across all stagnation temperature ratios considered. In addition to these insights, the study discusses the importance of proper thermal scaling and definition of adiabatic effectiveness when operating at high mainstream Mach number. It is demonstrated that the effects of high-speed flow challenge the efficacy of the conventional parameters used to characterize film cooling performance.

Abstract Cooling components in the hot section of a gas turbine is essential to component durability. Common methods of cooling include rib turbulators in internal passages and film cooling on external surfaces. The holes that produce the film cooling are fed from the internal channels often containing ribs. Consequently, there is an interdependence of internal heat transfer and external film cooling. The purpose of this study was to obtain a better understanding of the interaction of ribs and film cooling. To quantify the cooling performance the surface temperatures were measured from which overall effectiveness was calculated. For the experiments, additively manufactured test coupons were made of Inconel 718 to match engine Biot numbers. These test coupons had internal feed channels with and without ribs and had both cylindrical holes and meter diffuser shaped holes with 15° lateral expansion angles and a 1° forward expansion angle. A single rectangular channel was one type of feed channel. The other type of feed channels was individual circular channels with each circular channel supplying an individual film-cooling hole. The experimental results showed that the circular individual channels have 80% higher baseline overall effectiveness than the single rectangular channel without any film cooling. Ribbed turbulators without film cooling also increased the overall effectiveness by 21% for single rectangular channel and by 29% for the circular individual channels compared to the respective non-ribbed channels. While the film cooling increased the overall effectiveness of all geometries, the single rectangular channels had increased overall effectiveness levels by up to twice that of the no film-cooling case. On average the single rectangular channels had an 80% improvement from film cooling, whereas the individual channel feeds on average had only a 50% improvement given their high baseline effectiveness levels.

Abstract Gas turbine airfoils commonly use serpentine passages for internal cooling. To enhance the heat transfer in this internal passage, rib turbulators are often used. There have been many studies of rib turbulators to determine the optimum height, spacing, and angle; but almost all of these studies have been conducted using straight ducts. In this study, computational RANS predictions using the Realizable k-ϵ model and wind tunnel experimentation were used in conjunction to determine the optimum rib turbulator configuration for use in serpentine channels. The serpentine passage configuration had three U-bends and L/DH = 10 for each leg. The channel was placed below a simulated turbine airfoil wall which was flat, and the ribs in the channel were placed on this wall. Computations and experiments were completed with a smooth channel and with ribs at angles of −60°, 90°, and +60°, rib heights of e/H = 0.125 and 0.25, and spacing between ribs of P/e = 4 and 8. Studies were also done using turning vanes in the U-bends to reduce separation regions downstream of the U-bend. A key result from this study was that ribs oriented at −60° would produce similar pressure drops to the smooth serpentine channel. Flow field visualization showed that large separation regions after each U-bend of the serpentine channels were responsible for increased pressure drop, and the −60° ribs helped to redistribute the biased flow. Rib orientations of 90° and +60° experienced larger pressure drops than the −60° ribs both computationally and experimentally. Compared to the smooth channel, the channels with rib turbulators had only a small increase in heat transfer rates, ranging from 4% to 16% for varying rib angles. This small enhancement of heat transfer was attributed to the large heat transfer for the smooth serpentine channel caused by strong vortical flow generated downstream of the U-bends.

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° forward and lateral expansions (7-7-7 SI). Experiments without film cooling showed that the serpentine channel had 35% to 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 ϕ̿, 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 flow rate, the overall cooling effectiveness was ϕ̿=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.

Abstract Film cooling is used to protect turbine components from the extreme temperatures by ejecting coolant through arrays of holes to create an air buffer from the hot combustion gases. Limitations in traditional machining meant film cooling holes universally have sharp inlets, which create separation regions at the hole entrance. The present study uses experimental and computational data to show that these inlet separation are a major cause of performance variation in crossflow fed film cooling holes. Three-hole designs were experimentally tested by independently varying the coolant velocity ratio (VR) and the coolant channel velocity ratio (VRc) to isolate the effects of crossflow on hole performance. Leveraging additive manufacturing (AM) technologies, the addition of a 0.25D radius fillet to the inlet of a 7-7-7 shaped hole is shown to significantly improve diffuser usage and significantly reduce variation in performance with VRc. A second AM design used a very large radius of curvature inlet to reduce biasing caused by the inlet crossflow. Experiments showed that this “swept” hole design did minimize biasing of the coolant flow to one side of the shaped hole, and it significantly reduced variations due to varying VRc. RANS simulations at six VR and three VRc conditions were made for each geometry to better understand how the new geometries changed the velocity field within the hole. The sharp and rounded inlets were seen to have very similar tangential velocity fields and jet biasing. Both AM inlets created more uniform, slower velocity fields entering the diffuser. The results of this article indicate that large improvements in film cooling performance can be found by leveraging AM technology.