Adiabatic film cooling effectiveness measurements are obtained using pressure-sensitive paint (PSP) on a flat film cooled surface. The effects of blowing ratio and hole spacing are investigated for four multi-row arrays comprised of 8 rows containing 52 holes of 3.8 mm diameter with 20° inclination angles and hole length-to-diameter ratio of 11.2. The four arrays investigated have two different hole-to-hole spacings composed of cylindrical and diffuser holes. For the first case, lateral and streamwise pitches are 7.5 times the diameter. For the second case, pitch-to-diameter ratio is 14 in lateral direction and 10 in the streamwise direction. The holes are in a staggered arrangement. Adiabatic effectiveness measurements are taken for a blowing ratio range of 0.3 to 1.2 and a density ratio of 1.5, with CO2 injected as the coolant. A thorough boundary layer analysis is presented, and data was taken using hotwire anemometry with air injection, with boundary layer and turbulence measurements taken at multiple locations in order to characterize the boundary layer. Local effectiveness, laterally averaged effectiveness, boundary layer thickness, momentum thickness, turbulence intensity and turbulence length scale are presented. For the cylindrical holes, at the first row of injection, the film jets are still attached at a blowing ratio of 0.3. By a blowing ratio of 0.5, the jet is observed to lift off, and then impinge back onto the test surface. At a blowing ratio of 1.2, the jets lift off, but reattach much further downstream, spreading the coolant further along the test surface. A thorough uncertainty analysis has been conducted in order to fully understand the presented measurements and any shortcomings of the measurement technique. The maximum uncertainty of effectiveness and blowing ratio is 0.02 counts of effectiveness and 3 percent respectively.

Pressure-sensitive paint (PSP) can be a powerful tool in measuring the adiabatic film cooling effectiveness. There are two distinct sources of error for this measurement technique: the ability to experimentally obtain the data and the validity of the heat and mass transfer analogy for the problem being studied. This paper will assess the experimental aspect of this PSP measurement specifically for film cooling applications. Experiments are conducted in an effort to quantifiably bound expected errors associated with temperature nonuniformities in testing and photodegradation effects. Results show that if careful experimental procedures are put in place, both of these effects can be maintained to have less than 0.022 impact on effectiveness. Through accurate semi in situ calibration down to 4% atmospheric pressure, the near-hole distribution of effectiveness is measured with high accuracy. PSP calibrations are performed for multiple coupons, over multiple days. In addition, to reach a partial pressure of zero the calibration vessel was purged of all air by flowing CO2. The primary contribution of this paper lies in the uncertainty analysis performed on the PSP measurement technique. A thorough uncertainty analysis is conducted and described, in order to completely understand the presented measurements and any shortcomings of the PSP technique. This quantification results in larger, albeit more realistic, values of uncertainty and helps provide a better understanding of film cooling effectiveness measurements taken using the PSP technique. The presented uncertainty analysis takes into account all random error sources associated with sampling and calibration, from intensities to effectiveness. Adiabatic film cooling effectiveness measurements are obtained for a single row of film cooling holes inclined at 20 deg, with CO2 used as coolant. Data are obtained for six blowing ratios. Maps of uncertainty corresponding to each effectiveness profile are available for each test case. These maps show that the uncertainty varies spatially over the test surface and high effectiveness corresponds to low uncertainty. The noise floors can be as high as 0.04 at effectiveness levels of 0. Day-to-day repeatability is presented for each blowing ratio and shows that laterally averaged effectiveness data are repeatable within 0.02 effectiveness.

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.

Particle image velocimetry and particle scattering measurements are taken over a flat multi-row film cooled surface. The effect of coolant mass flow (characterized by blowing ratio, momentum flux ratio or velocity ratio) is investigated. The array investigated has 8 rows containing 52 holes of 3.8 mm diameter with inclination angles of 20° and hole length-to-diameter ratios of 11.2. The lateral and streamwise pitches are 7.5 times the diameter. The holes are in a staggered arrangement. PIV measurements are taken with CO2 as coolant injecting into an air primary stream. Two streamwise aligned planes are measured; a centerline plane and a plane spaced laterally half a diameter from centerline. The PIV results provide a quantitative view of the flow over eight rows of injection as well as the recovery region. Images are acquired at 8 Hz and have a spatial resolution of 10.3 interrogation windows per hole diameter. The tunnel has low freestream turbulence (1%) and is low speed, so the primary turbulence production mechanism is shearing. As a result, the M = 0.45 blowing ratio case shows higher levels of turbulent kinetic energy compared with the M = 1.05 blowing ratio case. Uncertainty in the PIV measurements is estimated via the correlation statistics method.

A flat plate film cooling experimental facility has enabled a study of the flowfield features and film performance characteristics of a single row of standard axial-shaped holes (ASH). Surface IR thermography and particle image velocimetry (PIV) were used to quantify blowing ratios impact on both flow and heat transfer in the ASH film row configuration. At constant freestream flow conditions, the blowing ratio was adjusted by altering the coolant flow rate. Performance was analyzed via IR measurements at three nominal blowing ratios by contours of adiabatic effectiveness as well as lateral and area-averaged adiabatic effectiveness. 2D PIV measurements were utilized as an additional flow diagnostic to examine the film jet structural variation with blowing ratio, and to provide a link between the behavior of the near-wall flowfield and the surface cooling.

The effectiveness of film cooling from short cooling holes, scaled to engine conditions, has been shown to be dependent on the nature of the internal coolant feed. A common method of supplying coolant to film cooling holes in engine components is through an internal cross-flow, which causes skewed effectiveness profiles on the surface of film cooled parts. For round axial holes, this effect causes coolant jets to more effectively spread across the surface. Additionally, for compound angle round holes, the direction of the cross-flow relative to the direction of injection has a substantial effect on film cooling effectiveness. A cross-flow directed counter to the span-wise direction of coolant injection has previously been shown to cause greater lateral jet spreading than cross-flow directed in-line with the span-wise injection direction. To better understand the phenomena responsible for the improved coolant spreading, two-dimensional thermal field profiles were measured downstream of compound angle film cooling holes fed by an internal cross-flow. A smooth-walled rectangular channel was used to produce an internal cross-flow in both a counter and in-line flow direction. Thermal field cross-section data was collected at three stream-wise locations: 0.7, 3.4, and 8.8 diameters downstream of the holes. Blowing ratios of 0.75 and 1.00 were studied at a density ratio of 1.5. Experiments were performed in a low speed recirculating wind tunnel at high mainstream turbulence with a thick approach boundary layer relative to the film cooling holes. It was found that the improved lateral spreading observed in the coolant jets fed by a counter cross-flow occurred due to the formation of a bulge on the downstream side of the jet.Copyright © 2015 by ASME

The use of impingement cooling to maintain acceptable material temperatures in cooled gas turbine components is one of the handful of very robust and widely employed cooling methods throughout propulsion and power-generating engines. Impingement cooling utilizes an available pressure differential and one or many appropriate nozzle confi gurations to generate high-speed jets that are directed at component interior surfaces for the purpose of cooling the walls. The potential design space for impingement cooling in these components is immense, yet the actual design space in use tends to be well characterized by a relatively few established correlations. Impingement cooling is used in virtually all of the hot gas path components of the engine, including the combustor, turbine vanes, blades, and shrouds, and even the rotating disks. The applications span jet Reynolds numbers from 10,000 to 500,000. The complexity of impingement cooling comes with the vast numbers of specifi c jet and surface geometry combinations that are involved with the actual components. Impingement cooling is a very fl exible method, and for this very reason must be made to conform to applications that are dictated by other design requirements, such as those from aerodynamics, mechanical structure, vibrations, low cycle fatigue, and creep rupture limits. This chapter presents the basics of gas turbine impingement cooling design, the various applications within the engine, and the typical limitations imposed upon its use. Generic use and application examples for single impingement jets, in-line rows of jets, and arrays of jets are described for the hot gas fl ow path components of gas turbines, including the combustor system and high-pressure turbine. Specifi c applications to airfoil aerodynamic leading and trailing edge regions, combustor liners, and rotating disks are provided, as well as emerging applications within confi ned channels, blade tips, and fi lm cooling. Signifi cant effects infl uencing impingement cooling www.witpress.com, ISSN 1755-8336 (on-line) WIT Transactions on State of the Art in Science and Engineering, Vol 76, © 2014 WIT Press doi:10.2495/978-1-84564-907-4/001 2 IMPINGEMENT JET COOLING IN GAS TURBINES are discussed including surface roughness, rotation, and accumulated cooling cross-fl ow degradation. The chapter wraps up with a discussion of the main limitations placed on impingement cooling due to available pressure, manufacturing, turbine operation, and system thermal management.

In gas turbine engines, film cooling holes are often fed by an internal cross-flow, with flow normal to the direction of the external flow around the airfoil. Many experimental studies have used a quiescent plenum to feed model film cooling holes and thus do not account for the effects of internal cross-flow. In this study, an experimental flat plate facility was constructed to study the effects of internal cross-flow on a row of cylindrical compound angle film cooling holes. Operating conditions were scaled, based on coolant hole Reynolds number and turbulence level, to match realistic turbine engine conditions. A cross-flow channel allowed for coolant to flow alternately in either direction perpendicular to the mainstream flow. Film cooling holes were operated at blowing ratios ranging from 0.5 to 2.0 at a density ratio of 1.5. There are relatively few studies available in literature that focus on the effects of cross-flow on film cooling performance, with no studies examining the effects of internal cross-flow on film cooling with round, compound angled holes. This study showed that significantly greater adiabatic effectiveness was achieved for cross-flow in the opposite direction of the span-wise direction of the coolant holes and provides possible explanations for this result.

Recent advances in computational power have made conjugate heat transfer simulations of fully conducting, film cooled turbine components feasible. However, experimental data available with which to validate conjugate heat transfer simulations is limited. This paper presents experimental measurements of external surface temperature on the suction side of a scaled up, fully conducting, cooled gas turbine vane. The experimental model utilizes the matched Bi method, which produces non-dimensional surface temperature measurements that are representative of engine conditions. Adiabatic effectiveness values were measured on an identical near adiabatic vane with an identical geometry and cooling configuration. In addition to providing a valuable data set for computational code validation, the data shows the effect of film cooling on the surface temperature of a film cooled part. As expected, in nearly all instances the addition of film cooling was seen to decrease the overall surface temperature. However, due to the effect of film injection causing early boundary layer transition, film cooling at a high momentum flux ratio was shown to actually increase surface temperature over 0.35 < s/C < 0.45.

This study compares surface pressure measurements and predictions for a high pressure turbine first-stage nozzle vane. The surface pressure measurements were taken in a 3D annular cascade, consisting of four airfoils and five passages. The cascade was uncooled, axisymmetric at both inner and outer endwalls, and reproduced the design intent Reynolds and Mach numbers of the real engine component. Static pressure measurements were taken along the airfoil profile at 15, 50, and 85% span, with duplicate midspan measurements across the four airfoils for assessing the tangential periodicity of the flow. Static pressure measurements were also taken on the inner and outer endwall surfaces of the center airfoil passage, with 40 measurement points uniformly distributed over each endwall. Three methods of surface pressure prediction were compared with the data: (1) a 2D inviscid CFD solution of a single airfoil passage at fixed spanwise locations, (2) a 3D RANS CFD solution of a single airfoil passage, and (3) a 3D RANS CFD solution of the full five-passage cascade domain. Both of the single-passage solutions assumed flowfield periodicity in the tangential direction and compared favorably to the center passage airfoil data. This finding suggested that the cascade center passage was sufficiently representative of the full-annulus turbomachine environment and validated the cascade for further experimental studies. The adjacent airfoil pressure measurements quantified the passage-to-passage variation in the cascade flowfield, and the 3D full-cascade CFD compared favorably with the peripheral airfoil data. The full-cascade CFD also compared favorably with the data on both endwalls: with an average and maximum deviation of 0.5 and 2%, respectively. These findings provide confidence in the 3D CFD methods for use in determining local flow rates from cooling/leakage geometry, and serve as an important first step toward validating the methods for real-engine blockage effects like coolant and endwall contouring.

This study investigated the interaction of thermal barrier coatings (TBC) and various film cooling configurations to provide a detailed assessment of the thermal protection on a first stage turbine vane. The internally cooled, scaled-up turbine vane used for this study was designed to properly model the conjugate heat transfer effects found in a real engine. The TBC material was selected to properly scale the thicknesses and thermal conductivities of the model to those of the engine. External surface temperatures, TBC-vane interface temperatures and internal temperatures were all measured over a range of internal coolant Reynolds numbers and mainstream turbulence intensities. The blowing ratio of the various film-cooling designs was also varied. The addition of TBC on the vane surface was found to increase the overall effectiveness of the vane surface just downstream of the coolant holes by up to 0.25 when no film cooling was present. The presence of the TBC significantly dampened the variations in overall effectiveness due to changes in blowing ratio which mitigated the detrimental effects of coolant jet separation. It was also discovered that with the presence of TBC standard round holes showed equivalent, if not better, performance when compared to round holes embedded in a shallow transverse trench.

Common gas turbine heat transfer analysis methods rely on the assumption that the driving temperature for heat transfer to a film cooled wall can be approximated by the adiabatic wall temperature. This assumption implies that the gas temperature above a film cooled adiabatic wall is representative of the overlying gas temperature on a film cooled conducting wall. This assumption has never been evaluated experimentally. In order for the adiabatic wall temperature as driving temperature for heat transfer assumption to be valid, the developing thermal boundary layer that exists above a conducting wall must not significantly affect the overriding gas temperature. In this paper, thermal fields above conducting and adiabatic walls of identical geometry and at the same experimental conditions were measured. These measurements allow for a direct comparison of the thermal fields above each wall in order to determine the validity of the adiabatic wall temperature as driving temperature for heat transfer assumption. In cases where the film cooling jet was detached, a very clear effect of the developing thermal boundary layer on the gas temperature above the wall was measured. In this case, the temperatures above the wall were clearly not well represented by the adiabatic wall temperature. For cases where the film cooling jet remained attached, differences in the thermal fields above the adiabatic and conducting wall were small, indicating a very thin thermal boundary layer existed beneath the coolant jet.

Conjugate heat transfer (CHT) calculation techniques for the heat transfer analysis of high-pressure turbines (HPT) have been developed during the past few years. Thus, it has become possible to take into account the coupling of the film, internal cooling, external gas flow and the metal diffusive heat transfer. The coupling problem may become extremely important in regions such as the airfoil leading edge and the vicinity of film hole breakout region where heat fluxes and thermal gradients are high. This article presents the results obtained using fully coupled 3D CHT simulations of a simplified film-cooled leading edge model and a NASA C3X vane with suction side film cooling. The results for the two cases are compared against experimental data obtained at University of Texas at Austin. The numerical simulations were conducted using the k-ω turbulence model. The leading edge model overall effectiveness predictions are in good agreement with the experiments, especially in the low blowing ratio range (1≤M≤2). For the C3X vane, the CHT results tend to underpredict the midspan and laterally averaged effectiveness due to film liftoff. However, the quantitative agreement is still reasonably good. The different levels of overall effectiveness agreement found between all cases are discussed.

Recent advances in computing power have made conjugate heat transfer simulations of turbine components increasingly popular; however, limited experimental data exists with which to evaluate these simulations. The primary parameter used to evaluate simulations is often the external surface temperature distribution, or overall effectiveness. In this paper, the overlying momentum and thermal boundary layers at various streamwise positions around a conducting, internally cooled simulated turbine vane were measured under low (Tu = 0.5%) and high (Tu = 20%) freestream turbulence conditions. Furthermore, experimental results were compared to computational predictions. In regions were a favorable pressure gradient existed, the thermal boundary layer was found to be significantly thicker than the accompanying momentum boundary layer. Elevated freestream turbulence had the effect of thickening the thermal boundary layer much more effectively than the momentum boundary layer over the entire vane. This data is valuable in understanding the conjugate heat transfer effects on the vane as well as serving as a tool for computational code evaluation.

An experimental and computational conjugate heat transfer study of an internally cooled, scaled-up simulated turbine vane with internal rib turbulators was performed. The conjugate nature of the model allowed for the effects of the internal ribs to be seen on the external overall effectiveness distribution. The enhanced internal heat transfer coefficient caused by the ribs increased the cooling capacity of the internal cooling circuit, lowering the overall metal temperature. External surface temperatures, internal surface temperatures, and coolant inlet and exit temperatures were measured and compared to data obtained from a non-ribbed model over a range of internal coolant Reynolds numbers. Internal rib turbulators were found to increase the overall effectiveness on the vane external surface by up to 50% relative to the non-ribbed model. Additionally, comparisons between the experimental measurements and computational predictions are presented.

Heat transfer coefficients were measured downstream of a row of shaped film cooling holes, as well as elliptical, diffuser, and teardrop shaped dimples, simulating depressions due to film coolant holes of different shapes. These features were placed on the suction side of a simulated gas turbine vane. The dimples were used as approximations to film cooling holes after the heat transfer levels downstream of active fan shaped film cooling holes was found to be independent of film cooling. The effects of the dimples were tested with varying approach boundary layers, freestream turbulence intensity, and Reynolds numbers. For the case of an untripped (transitional) approach boundary layer, all dimple shapes caused approximately a factor of 2 increase in heat transfer coefficient relative to the smooth baseline condition due to the dimples effectively causing boundary layer transition downstream. The exact augmentation varied depending on the dimple geometry: diffuser shapes causing the largest augmentation and teardrop shapes causing the lowest augmentation. For tripped (turbulent boundary layer) approach conditions, the dimple shapes all caused the same 20% augmentation relative to the smooth tripped baseline. The already turbulent nature of the tripped approach flow reduces the effect that the dimples have on the downstream heat transfer coefficient.

Skin friction coefficients and heat transfer coefficients are measured for a range of regular and random roughnesses on the suction side of a simulated gas turbine vane. The skin friction coefficients are calculated using boundary layer data and the momentum integral method. High resolution surface temperature data measured with an IR camera yield local heat transfer values. 80 grit, 50 grit, 36 grit, and 20 grit sandpapers along with a regular array of conical roughness elements are tested. Measured skin friction coefficient data show that the conical roughness array behaves very similar to the 50 grit, 36 grit, and 20 grit sandpapers in terms of the effect of the roughness on the hydrodynamic boundary layer. In terms of heat transfer, the conical roughness array is most similar to the 80 grit sandpaper, which are both lower than the roughest sandpapers tested. These data show that the particular regular array of roughness elements tested has fundamentally different behavior than randomly rough surfaces for this position on the simulated turbine vane. In addition, this difference is in the opposite direction as seen in previous experimental studies. In order to draw a more general conclusion about the nature of random and regular roughness, a parametric study of regular roughness arrays should be performed.

Film cooling adiabatic effectiveness and heat transfer coefficients for cylindrical holes embedded in a 1d transverse trench on the suction side of a simulated turbine vane were investigated to determine the net heat flux reduction. For reference, measurements were also conducted with standard inclined, cylindrical holes. Heat transfer coefficients were determined with and without upstream heating to isolate the hydrodynamic effects of the trench and to investigate the effects of the thermal approach boundary layer. Also the effects of a tripped versus an un-tripped boundary layer were explored. For both the cylindrical holes and the trench, heat transfer augmentation was much greater with no tripping of the approach flow. A further increase in heat transfer augmentation was caused by use of upstream heating, with as much as a 150% augmentation with the trench. With a tripped approach flow the heat transfer augmentation was much less. The net heat flux reduction for the trench was found to be significantly higher than for the row of cylindrical holes.