Most studies of turbine airfoil film cooling in laboratory test facilities have used relatively large plenums to feed flow into the coolant holes. However, a more realistic inlet condition for the film cooling holes is a relatively small channel. Previous studies have shown that the film cooling performance is significantly degraded when fed by perpendicular internal crossflow in a smooth channel. In this study, angled rib turbulators were installed in two geometric configurations inside the internal crossflow channel, at 45 deg and 135 deg, to assess the impact on film cooling effectiveness. Film cooling hole inlets were positioned in both prerib and postrib locations to test the effect of hole inlet position on film cooling performance. A test was performed independently varying channel velocity ratio and jet to mainstream velocity ratio. These results were compared to the film cooling performance of previously measured shaped holes fed by a smooth internal channel. The film cooling hole discharge coefficients and channel friction factors were also measured for both rib configurations with varying channel and inlet velocity ratios. Spatially averaged film cooling effectiveness is largely similar to the holes fed by the smooth internal crossflow channel, but hole-to-hole variation due to inlet position was observed.

Internal crossflow, or internal flow that is perpendicular to the overflowing mainstream, reduces film cooling effectiveness by disrupting the diffusion of coolant at the exit of axial shaped holes. Previous experimental investigations have shown that internal crossflow causes the coolant to bias toward one side of the diffuser and that the severity of the biasing scales with the inlet velocity ratio, VR i , or the ratio of crossflow velocity to the jet velocity in the metering section of the hole. It has been hypothesized and computationally predicted that internal crossflow produces an asymmetric swirling flow within the hole that causes the coolant to bias in the diffuser and that biasing contributes to ingestion of hot mainstream gas into the hole, which is undesirable. However, there are no experimental measurements as of yet to confirm these predictions. In the present study, in- and near-hole flow field and thermal field measurements were performed to investigate the flow structures and mainstream ingestion for a standard axial shaped hole fed by internal crossflow. Three different inlet velocity ratios of VR i = 0.24, 0.36, and 0.71 were tested at varying injection rates. Measurements were made in planes normal to the nominal direction of coolant flow at the outlet plane of the hole and at two downstream locations—x/d = 0 and 5. The predicted swirling structure was observed for the highest inlet velocity ratio and flow within the hole was shown to scale with VR i . Ingestion within the diffuser was significant and also scaled with VR i . Downstream flow and thermal fields showed that increased biasing contributed to more severe jet detachment and coolant dispersion away from the surface.

Film cooling can have a significant effect on the heat transfer coefficient (HTC) between the overflowing freestream gas and the underlying surface. This study investigated the influence of approach flow characteristics, including the boundary layer thickness and character (laminar and turbulent), as well as the approach flow Reynolds number, on the HTC. The figure of merit for this study was the HTC augmentation, that is, the ratio of HTCs for a cooled versus uncooled surface. A heated foil surface provided a known heat flux, allowing direct measurement of HTC and augmentation. The foil was placed both upstream and downstream of the film cooling holes, in order to generate an approaching thermal boundary layer, as representative of actual engine conditions. High-resolution IR thermography provided spatially resolved HTC augmentation data. An open-literature shaped-hole design was used, known as the 7-7-7 hole, in order to compare with existing results in the literature. A variety of blowing conditions were tested from M = 0.5 to 3.0. Two elevated density ratios of DR = 1.20 and DR = 1.80 were used. The results indicated that turbulent boundary layer thickness had a modest effect on HTC augmentation, whereas a very high level of augmentation was observed for a laminar approach boundary layer. The presence of upstream heating greatly increased the HTC augmentation in the near-hole region, although these effects died out by 10–15 diameters from the holes.

In gas turbine engines, film cooling holes are commonly fed with an internal crossflow, the magnitude of which has been shown to have a notable effect on film cooling effectiveness. In Part I of this study, as well as in a few previous studies, the magnitude of internal crossflow velocity was shown to have a substantial effect on film cooling effectiveness of axial shaped holes. There is, however, almost no data available in the literature that shows how internal crossflow affects compound angle shaped film cooling holes. In Part II, film cooling effectiveness, heat transfer coefficient augmentation, and discharge coefficients were measured for a single row of compound angle shaped film cooling holes fed by internal crossflow flowing both in-line and counter to the spanwise direction of coolant injection. The crossflow-to-mainstream velocity ratio was varied from 0.2 to 0.6 and the injection velocity ratio was varied from 0.2 to 1.7. It was found that increasing the magnitude of the crossflow velocity generally caused degradation of the film cooling effectiveness, especially for in-line crossflow. An analysis of jet characteristic parameters demonstrated the importance of crossflow effects relative to the effect of varying the film cooling injection rate. Heat transfer coefficient augmentation was found to be primarily dependent on injection rate, although for in-line crossflow, increasing crossflow velocity significantly increased augmentation for certain conditions.

The effect of feeding shaped film cooling holes with an internal crossflow is not well understood. Previous studies have shown that internal crossflow reduces film cooling effectiveness from axial shaped holes, but little is known about the mechanisms governing this effect. It was recently shown that the crossflow-to-mainstream velocity ratio is important, but only a few of these crossflow velocity ratios have been studied. This effect is of concern because gas turbine blades typically feature internal passages that feed film cooling holes in this manner. In this study, film cooling effectiveness was measured for a single row of axial shaped cooling holes fed by an internal crossflow with crossflow-to-mainstream velocity ratio varying from 0.2 to 0.6 and jet-to-mainstream velocity ratios varying from 0.3 to 1.7. Experiments were conducted in a low speed flat plate facility at coolant-to-mainstream density ratios of 1.2 and 1.8. It was found that film cooling effectiveness was highly sensitive to crossflow velocity at higher injection rates while it was much less sensitive at lower injection rates. Analysis of the jet shape and lateral spreading found that certain jet characteristic parameters scale well with the crossflow-to-coolant jet velocity ratio, demonstrating that the crossflow effect is governed by how coolant enters the film cooling holes.

Knowing the heat transfer coefficient augmentation is imperative to predicting film cooling performance on turbine components. In the past, heat transfer coefficient augmentation was generally measured at unit density ratio to keep measurements simple and uncertainty low. Some researchers have measured heat transfer coefficient augmentation while taking density ratio effects into account, but none have made direct temperature measurements of the wall and adiabatic wall to calculate h f /h 0 at higher density ratios. This work presents results from measuring the heat transfer coefficient augmentation downstream of shaped holes with a 7 deg forward and lateral expansion at DR = 1.0, 1.2, and 1.5 on a flat plate using a constant heat flux surface. The results showed that the heat transfer coefficient augmentation was low while the jets were attached to the surface and increased when the jets started to separate. At DR = 1.0, h f /h 0 was higher for a given blowing ratio than at DR = 1.2 and DR = 1.5. However, when velocity ratios are matched, better correspondence was found at the different density ratios. Surface contours of h f /h 0 showed that the heat transfer was initially increased along the centerline of the jet, but was reduced along the centerline at distances farther downstream. The decrease along the centerline may be due to counter-rotating vortices sweeping warm air next to the heat flux plate toward the center of the jet, where they sweep upward and thicken the thermal boundary layer. This warming of the core of the coolant jet over the heated surface was confirmed with thermal field measurements.

Facilities such as the Turbine Research Facility (TRF) at the Air Force Research Laboratory have been acquiring uncooled heat transfer measurements on full-scale metallic airfoils for several years. The addition of cooling flow to this type of facility has provided new capabilities and new challenges. Two primary challenges for cooled rotating hardware are that the true local film temperature is unknown, and cooled thin-walled metallic airfoils prohibit semi-infinite heat conduction calculation. Extracting true local adiabatic effectiveness and the heat transfer coefficient from measurements of surface temperature and surface heat transfer is therefore difficult. In contrast, another cooling parameter, the overall effectiveness ( ϕ ), is readily obtained from the measurements of surface temperature, internal coolant temperature, and mainstream temperature. The overall effectiveness is a normalized measure of surface temperatures expected for actual operating conditions and is thus an important parameter that drives the life expectancy of a turbine component. Another issue is that scaling ϕ from experimental conditions to engine conditions is dependent on the heat transfer through the part. It has been well-established that the Biot number must be matched for the experimentally measured ϕ to match ϕ at engine conditions. However, the thermal conductivity of both the metal blade and the thermal barrier coating changes substantially from low-temperature to high-temperature engine conditions and usually not in the same proportion. This paper describes a novel method of replicating the correct thermal behavior of the thermal barrier coating (TBC) relative to the metal turbine while obtaining surface temperature measurements and heat fluxes. Furthermore, this paper describes how the ϕ value obtained at the low-temperature conditions can be adjusted to predict ϕ at high-temperature engine conditions when it is impossible to match the Biot number perfectly.

Early stage gas turbine blades feature complicated internal geometries in order to enhance internal heat transfer and to supply coolant for film cooling. Most film cooling experiments decouple the effect of internal coolant feed from external film cooling effectiveness, even though engine parts are commonly fed by cross-flow and feature internal rib turbulators which can affect film cooling. Experiments measuring adiabatic effectiveness were conducted to investigate the effects of turbulated perpendicular cross-flow on a row of 45 deg compound angle cylindrical film cooling holes for a total of eight internal rib configurations. The ribs were angled to the direction of prevailing internal cross-flow at two different angles: 45 deg or 135 deg. The ribs were also positioned at two different spanwise locations relative to the cooling holes: in the middle of the cooling hole pitch and slightly intersecting the holes. Experiments were conducted at a density ratio of DR = 1.5 for a range of blowing ratios including M = 0.5, 0.75, 1.0, 1.5, and 2.0. This study demonstrates that peak effectiveness can be attained through the optimization of cross-flow direction relative to the compound angle direction and rib configuration, verifying the importance of hole inlet conditions in film cooling experiments. It was found that ribs tend to reduce adiabatic effectiveness relative to a baseline, smooth-walled configuration. Rib configurations that directed the internal coolant forward in the direction of the mainstream resulted in higher peak adiabatic effectiveness. However, no other parameters could consistently be identified correlating to increased film cooling performance. It is likely that a combination of factors is responsible for influencing performance, including internal local pressure caused by the ribs, the internal channel flow field, in-hole vortices, and jet exit velocity profiles. This study also attempted to replicate the possibility that film cooling holes may intersect ribs and found that a hole which partially intersects a rib still maintains moderate levels of effectiveness.

Film cooling designs are often evaluated experimentally and characterized in terms of their spatial distributions of adiabatic effectiveness, η , which is the nondimensionalized form of the adiabatic wall temperature, T aw . Additionally, film cooling may alter the convective heat transfer coefficient with the possibility of an increase in h that offsets the benefits of reduced T aw . It is therefore necessary to combine these two effects to give some measure of the benefit of film cooling. The most frequently used method is the net heat flux reduction (NHFR), which gives the fractional reduction in heat flux that accompanies film cooling for the hypothetical case of constant wall temperature. NHFR is imperfect in part due to the fact that this assumption does not account for the primary purpose of film cooling—to reduce the metal temperature to an acceptable level. In the present work, we present an alternative method of evaluating film cooling performance that yields the reduction in metal temperature, or in the nondimensional sense, an increase in ϕ that would be predicted with film cooling. This Δ ϕ approach is then applied using experimentally obtained η and h/h 0 values on a simulated turbine blade leading edge region. The delta-phi approach agrees well with the legacy NHFR technique in terms of the binary question of whether the film cooling is beneficial or detrimental, but provides greater insight into the temperature reduction that a film cooling design would provide an actual turbine component. For example, instead of giving an area-averaged NHFR = 0.67 (indicating a 67% reduction in heat flux through film cooling) on the leading edge region with M = 0.5, the Δ ϕ approach indicates an increase in ϕ of 0.061 (or a 61 K surface temperature decrease with a notional value of T ∞ −T c = 1000 K). Alternatively, the technique may be applied to predict the maximum allowable increase in T ∞ against which a film cooling scheme could protect.

Two-dimensional thermal profiles were experimentally measured downstream of a single row of film cooling holes on both an adiabatic and a matched Biot number model turbine vane. The measurements were taken as a comparison to computational simulations of the same model and flow conditions. Previously, adiabatic and overall effectiveness comparisons have been made between experimental and computational data. To improve computational models of the evolution of a film cooling jet as it propagates downstream, the thermal field above the vane, not just the footprint on the vane surface, must be analyzed. This study expands these data to include 2D thermal fields above the vane at 0, 5, and 10 hole diameters downstream of the film cooling holes. Four blowing ratios were tested, M = 0.28, 0.65, 1.11, and 2.41. In each case, the computational jets remained colder than the experimental jets because they did not diffuse into the mainstream as quickly. In addition, the computational results for the higher two blowing ratios exhibited the effects of the kidney vortex commonly studied in film cooling, but the experimental thermal fields were not dominated by this vortex. Finally, in comparing results above adiabatic and matched Biot number models, these thermal fields allow for an accurate analysis of whether or not the adiabatic wall temperature was a reasonable estimate of the driving temperature for heat transfer.