In this study, the effect of shape variables on the film cooling effectiveness of the compound angled asymmetric laidback fan shaped hole was experimentally investigated, and the optimum values of select design variables were presented. Among the shape variables of the compound angled asymmetric laidback fan shaped hole, the windward and leeward lateral expansion angles and the compound angle were selected as design variables. Test points were chosen using the central composite design method, and the selected design variables were optimized using the Kriging model. The film cooling effectiveness was measured using the PSP technique, and the experiment was conducted under the two density ratios of 1.5 and 2.0 and four blowing ratios of 1.0, 1.5, 2.0, and 2.5. Experimental results showed that the film cooling performance was improved for higher density ratios than lower density ratios. The main effects analysis indicated that larger windward and leeward lateral expansion angles induced higher film cooling effectiveness; however, the compound angle did not show consistent results. For the optimized hole at the density ratio 2.0, the results indicated that the overall averaged film cooling effectiveness of the optimized compound angled asymmetric laidback fan shaped hole was higher than that of the optimized fan shape holes of previous literature.

An experimental study was performed to investigate the effects of the arrangement of fan-shaped film cooling holes and density ratio (DR) on heat transfer coefficient augmentation. Both single- and multi-row fan-shaped film cooling holes were considered. For the multi-row fan-shaped holes, the heat transfer coefficient was measured at DRs of 1 and 2, and both staggered and inline arrangements of holes were considered. For the single-row fan-shaped holes, DR = 1.0, 1.5, 2.0, and 2.5 and M = 1.0 and 2.0 conditions were tested. The mainstream velocity was 20 m/s, and the turbulence intensity and boundary layer thickness were 3.6% and 6 mm, respectively. The heat transfer coefficient was measured using the one-dimensional transient infrared thermography method. The results show that an increased heat transfer coefficient augmentation is observed between film cooling holes for the case with a smaller hole pitch and higher blowing ratio. For the given fan-shaped hole parameters, the effects of the row-to-row distance and hole arrangement are not significant. In addition, as the velocity difference between the mainstream and coolant increases, the heat transfer coefficient ratio increases.

In this study, the effect of mainstream velocity on the optimization of a fan-shaped hole on a flat plate was experimentally investigated. The experiment was conducted by changing the forward expansion angle (${\mathsf{\beta }}_{\mathrm{fwd}}$), lateral expansion angle (${\mathsf{\beta }}_{\mathrm{lat}}$), and metering length ratio (${\mathrm{L}}_{\mathrm{m}}/\mathrm{D}$) of the film-cooling hole. A total of 13 cases extracted using the Box–Behnken method were considered to examine the effect of the shape parameters of the film-cooling hole under a 90 m/s mainstream velocity condition, and the results were compared with the results derived under a mainstream velocity of 20 m/s. One density ratio (DR = 2.0) and a blowing ratio (M) ranging from 1.0 to 2.5 were considered, and the pressure-sensitive paint (PSP) technique was applied for the film-cooling effectiveness (FCE). As a result of the experiment, the optimized hole showed a 49.3% improvement in the overall averaged FCE compared to the reference hole with DR = 2.0 and M = 2.0. As the blowing ratio increased, the hole exit area tended to increase, and this tendency was the same as that in the 20 m/s mainstream condition.

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.

In the vicinity of gas turbine blades, a complex flow field is formed due to the flow separation, reattachment, and secondary flows, and this results in a locally non-uniform and high heat transfer on the surfaces. The present study experimentally investigates the effects of leakage flow through the slot between the gas turbine vane and blade rows on the film cooling effectiveness of the forward region of the shroud ring segment. The experiment is carried out in a linear cascade with five blades. Instead of the vane, a row of rods at the location of the vane trailing edge is installed to consider the wake effect. The leakage flow is introduced through the slot between the vane and blade rows, and additional coolant air is injected from the cooling holes installed at the vane’s outer zone. The effects of the slot geometry, cooling hole configuration, and blowing ratio on the film cooling effectiveness are experimentally investigated using the pressure-sensitive paint (PSP) technique. CO₂ gas and a mixture of SF₆ and N₂ (25% + 75%) are used to simulate the leakage flow to the mainstream density ratios of 1.5 and 2.0, respectively. The results indicate that the area-averaged film cooling effectiveness is affected more by the slot width than by the cooling hole configuration at the same injection conditions, and the lower density ratio cases show higher film cooling effectiveness than the higher density ratio case at the same cooling configuration.

In this study, a comparison of aerodynamic loss of gas turbine vane with various trailing edge configurations was experimentally performed. For the trailing edge, three cooling schemes (the cutback slot, the central injection slot, and the pressure side hole) and three levels of trailing edge diameter (7.5, 10.0, and 12.5% of the throat width, respectively) were considered. The total pressure loss was measured in a five-vane linear cascade at the low-speed mainstream condition (20 m/s) using a Kiel probe and a two-axis traversing device. The measurement region was constrained to the near mid-span region, and the measurement step size was 1 mm. The results showed that the distribution of the total pressure loss was strongly affected by cooling schemes. The vane with the pressure side hole showed a similar distribution to that of the baseline in which a cooling scheme was not applied. It was also shown that the total pressure loss increased as the trailing edge diameter increased regardless of the cooling schemes. Except for the cutback slot cases, there was an inflection point at which the tendency of the area-averaged total pressure loss to the blowing ratio was changed. Through the present study, correlation equations for each cooling scheme were derived against the trailing edge diameter and the blowing ratio.