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General Interest

2006;():1-14. doi:10.1115/GT2006-90002.

The advancement of turbine cooling has allowed engine design to exceed normal material temperature limits, but it has introduced complexities that have accentuated the thermal issues greatly. Cooled component design has consistently trended in the direction of higher heat loads, higher through-wall thermal gradients, and higher in-plane thermal gradients. The present discussion seeks to identify ten major thermal issues, or opportunities, that remain for the turbine hot gas path today. These thermal challenges are commonly known in their broadest forms, but some tend to be little discussed in a direct manner relevant to gas turbines. These include uniformity of internal cooling, ultimate film cooling, micro cooling, reduced incident heat flux, secondary flows as prime cooling, contoured gas paths, thermal stress reduction, controlled cooling, low emission combustor-turbine systems, and regenerative cooling. Evolutionary or revolutionary advancements concerning these issues will ultimately be required in realizable engineering forms for gas turbines to breakthrough to new levels of performance. Herein lies the challenge to researchers and designers. It is the intention of this summary to provide a concise review of these issues, and some of the recent solution directions, as an initial guide and stimulation to further research.

Commentary by Dr. Valentin Fuster
2006;():15-24. doi:10.1115/GT2006-90006.

An experimental and numerical study was conducted to determine the thermal performance of V-shaped ribs in a rectangular channel with an aspect ratio of 2:1. Local heat transfer coefficients were measured using the steady state thermochromic liquid crystal technique. Periodic pressure losses were obtained with pressure taps along the smooth channel sidewall. Reynolds numbers from 95,000 to 500,000 were investigated with V-shaped ribs located on one side or on both sides of the test channel. The rib height-to-hydraulic diameter ratios (e/Dh ) were 0.0625 and 0.02, and the rib pitch-to-height ratio (P/e) was 10. In addition, all test cases were investigated numerically. The commercial software FLUENT™ was used with a two-layer k-ε turbulence model. Numerically and experimentally obtained data were compared. It was determined that the heat transfer enhancement based on the heat transfer of a smooth wall levels off for Reynolds numbers over 200,000. The introduction of a second ribbed sidewall slightly increased the heat transfer enhancement whereas the pressure penalty was approximately doubled. Diminishing the rib height at high Reynolds numbers had the disadvantage of a slightly decreased heat transfer enhancement, but benefits in a significantly reduced pressure loss. At high Reynolds numbers small-scale ribs in a one-sided ribbed channel were shown to have the best thermal performance.

Commentary by Dr. Valentin Fuster
2006;():25-35. doi:10.1115/GT2006-90021.

Fan-shaped film-cooling holes have been shown to provide superior cooling performance to cylindrical holes along flat-plates and turbine airfoils over a large range of different conditions. Benefits of fan-shaped holes include less required cooling air for the same performance, increased part lifetime, and fewer required holes. The major drawback however, is increased manufacturing cost and manufacturing difficulty, particularly for the vane platform region. To this point, there have only been extremely limited comparisons between cylindrical and shaped holes on a turbine endwall at either low or high freestream turbulence conditions. This study presents film-cooling effectiveness measurements on an endwall surface in a large-scale, low-speed, two-passage, linear vane cascade. Results showed that film-cooling effectiveness decreased with increasing blowing rate for the cylindrical holes, indicating jet lift-off. However, the fan-shaped passage showed increased film-cooling effectiveness with increasing blowing ratio. Overall, fan-shaped holes increased film-cooling effectiveness by an average of 75% over cylindrical holes for constant cooling flow.

Topics: Cooling , Turbulence
Commentary by Dr. Valentin Fuster
2006;():37-47. doi:10.1115/GT2006-90034.

Film cooling effectiveness measurements under rotation were performed on the rotor blade platform using a pressure sensitive paint (PSP) technique. The present study examines, in particular, the film cooling effectiveness due to purging of coolant from the wheel-space cavity through the circumferential clearance gap provided between the stationary and rotating components of the turbine. The experimental investigation is carried out in a new three-stage turbine facility, recently designed and taken into operation at the Turbomachinery Performance and Flow Research Laboratory (TPFL) of Texas A&M University. This new turbine rotor has been used to facilitate coolant injection through this stator-rotor gap upstream of the 1st stage rotor blade. The gap was inclined at 25° to mainstream flow to allow the injected coolant to form a film along the passage platform. The effects of turbine rotating conditions on the blade platform film cooling effectiveness were investigated at three speeds of 2550 rpm, 2000 rpm and 1500 rpm with corresponding incidence angles of 23.2°, 43.4° and 54.8° respectively. Four different coolant-to-mainstream mass flow ratios varying from 0.5% to 2.0% were tested at each rotational speed. Aerodynamic measurements were performed at the 1st stage stator exit using a radially traversed five-hole probe to quantify the mainstream flow at this station. Results indicate that film cooling effectiveness increases with an increase in the coolant-to-mainstream mass flow ratios for all turbine speeds. Higher turbine rotation speeds show more local film cooling effectiveness spread on the platform with increasing magnitudes.

Commentary by Dr. Valentin Fuster
2006;():49-60. doi:10.1115/GT2006-90048.

High-pressure turbine blade tips operate in a highly complex flow environment that makes designing new blades for increased life difficult. Computational fluid dynamics simulations of the tip flow field may be able to guide new designs to improve the blade life, but the analysis techniques need to be verified against detailed measurements before they can be applied. The current paper presents measurements of heat flux and pressure in the blade tip region of a modern one-and-one-half stage high-pressure turbine operating at design corrected conditions in a rotating rig. Both flat tip and recessed, or squealer, tip blades were used in the experiments. The measurements indicate that the recessed tip, used in the majority of modern turbines to minimize blade damage from rubs, increases the blade heat load overall, and creates several hot spots on the floor of the recess for an uncooled airfoil. The tip data also showed there were significant unsteady variations in the heat load at the vane passing frequency. Steady state CFD calculations were completed for both flat and squealer tip configurations to examine if the analysis could capture the details that were measured. The CFD, while not capable of estimating the unsteady heat load component and generally over predicting the overall heat flux by 10–25%, did capture the measured heat flux trends in the recessed tip. These results show that steady-state CFD analysis can be useful in predicting the complex flow field and heat load distribution in turbine blade tips to help guide future blade designs.

Commentary by Dr. Valentin Fuster
2006;():61-69. doi:10.1115/GT2006-90051.

A technical assessment of the potential improvements resulting from the application of active control of film cooling is discussed in this paper. Two methods of film-cooling jet excitation are theoretically evaluated. Film-cooling control through pulsation of the coolant stream mass flow rate is further examined on the basis of existing experimental and preliminary computational results. A potentially optimal range of coolant pulsation parameters are identified through theoretical and mechanistic arguments based on understanding of basic physical phenomena occurring in pulsed film-cooling flow, i.e. that of a pulsed jet in cross-flow. Guidelines for the choice of potentially optimum frequency and duty cycle windows of the controlling flow pulsations are identified, as well as the related mean and peak-to-peak blowing ratios. Evidence from existing experimental data and computations are invoked to validate these guidelines. Because of the limited experimental/computational data on this subject further work is necessary in order to fully verify the aforementioned flow-pulsing parameter ranges.

Commentary by Dr. Valentin Fuster
2006;():71-80. doi:10.1115/GT2006-90052.

The paper presents an experimental study of heat/mass transfer coefficient in 4:1 aspect ratio smooth channels with non-uniform cross-sections. Curved leading and trailing edges are studied, for two curvatures of 9.06 m−1 (0.23 in−1 ) and 15.11 m−1 (0.384 in−1 ) and for two different curvature configurations. One configuration has curved walls with curvature corresponding to the blade profile (positive curvature on both leading and trailing walls), and the other configuration has leading and trailing walls that curve inwards into the coolant passage (negative curvature on the leading surface and positive curvature on the trailing surface). A detailed study at Re = 10,000 with rotation numbers in the range of 0–0.07 is undertaken for the two different curvature configurations. All experiments are done for a 90° passage-orientation with respect to the plane of rotation. The experiments are conducted in a rotating two-pass coolant channel facility using the naphthalene sublimation technique. Only the radially outward flow is considered for the present study. The span-wise mass transfer distributions of fully developed regions of the channel walls are also presented. The mass transfer data from the curved wall channels is compared to those from a smooth 4:1 rectangular duct with similar flow parameters. The local mass transfer data is analyzed mainly for the fully developed region, and area-averaged results are presented to delineate the effect of the rotation number. Heat transfer enhancement especially in the leading wall is seen for the lower curvature channels, and there is a subsequent reduction in the higher curvature channel, when compared to the 4:1 rectangular smooth channel. This indicates that an optimal channel wall curvature exists for which heat transfer is the highest.

Commentary by Dr. Valentin Fuster
2006;():81-90. doi:10.1115/GT2006-90067.

Gas turbines are often subjected to conditions where dirt and sand are ingested into the engine during takeoffs and landings. Given most aero engines do not have filtration systems, particulates can be present in both the main gas path and coolant streams. Particulates can block coolant passages and film-cooling holes that lead to increased airfoil temperatures caused by reduced coolant available for a given pressure ratio across the cooling holes. This study investigated the effects of sand blockage on film-cooling holes placed in a leading edge coupon. The coupon was tested to determine the reduction in flow parameter for a range of pressure ratios, coolant temperatures, metal temperatures, number of cooling holes, sand amounts, and sand diameters. Depending upon conditions, blockages characterized by reduced coolant flow can be as high as 10%.

Topics: Cooling , Sands
Commentary by Dr. Valentin Fuster
2006;():91-102. doi:10.1115/GT2006-90089.

To protect hot turbine components, cooler air is bled from the high pressure section of the compressor and routed around the combustor where it is then injected through the turbine surfaces. Some of this high pressure air also leaks through the mating gaps formed between assembled turbine components where these components experience expansions and contractions as the turbine goes through operational cycles. This study presents endwall adiabatic effectiveness levels measured using a scaled up, two-passage turbine vane cascade. The focus of this study is evaluating the effects of thermal expansion and contraction for the combustor-turbine interface. Increasing the mass flow rate for the slot leakage between the combustor and turbine showed increased local adiabatic effectiveness levels while increasing the momentum flux ratio for the slot leakage dictated the coverage area for the cooling. With the mass flow held constant, decreasing the combustor-turbine interface width caused an increase in uniformity of coolant exiting the slot, particularly across the pressure side endwall surface. Increasing the width of the interface had the opposite effect thereby reducing coolant coverage on the endwall surface.

Commentary by Dr. Valentin Fuster
2006;():103-116. doi:10.1115/GT2006-90108.

This paper presents the first experimental and numerical work of film effectiveness performance for a novel film cooling method with an arrowhead-shaped hole geometry. Experimental results demonstrate that the proposed hole geometry improves the film effectiveness on both suction and pressure surface of a generic turbine airfoil. Film effectiveness data for a row of the holes are compared with that of fan-shaped holes at the same inclination angle of 35° to the surface on a large-scale airfoil model at engine representative Reynolds number and Mach number in a high speed tunnel with moderately elevated temperature mainstream flow. The film effectiveness data are collected using pressure sensitive paint (PSP). Numerical results show that the coolant film with the proposed hole geometry remains well attached to the surface and diffuses in the lateral direction in comparison with the conventional laidback fan-shaped holes for coolant to mainstream blowing ratios of 0.6 to 3.5.

Topics: Cooling , Geometry
Commentary by Dr. Valentin Fuster
2006;():117-126. doi:10.1115/GT2006-90153.

This paper describes an experimental study of heat transfer in a radially rotating square duct with two opposite walls roughened by 45° staggered ribs. Air coolant flows radially outward in the test channel with experiments to be undertaken that match the actual engine conditions. Laboratory-scale heat transfer measurements along centerlines of two rib-roughened surfaces are performed with Reynolds number (Re), rotation number (Ro) and density ratio (Δρ/ρ) in the ranges of 7500–15000, 0–1.8 and 0.076–0.294. The experimental rig permits the heat transfer study with the rotation number considerably higher than those studied in other researches to date. The rotational influences on cooling performance of the rib-roughened channel due to Coriolis forces and rotating buoyancy are studied. A selection of experimental data illustrates the individual and interactive impacts of Re, Ro and buoyancy number on local heat transfer. A number of experimental-based observations reveal that the Coriolis force and rotating buoyancy interact to modify heat transfer even if the rib induced secondary flows persist in the rotating channel. Local heat transfer ratios between rotating and static channels along the centerlines of stable and unstable rib-roughened surfaces with Ro varying from 0.1 to 1.8 are in the ranges of 0.6–1.6 and 1–2.2 respectively. Empirical correlations for periodic flow regions are developed to permit the evaluation of interactive and individual effects of rib-flows, convective inertial force, Coriolis force and rotating buoyancy on heat transfer.

Topics: Heat transfer , Ducts
Commentary by Dr. Valentin Fuster
2006;():127-137. doi:10.1115/GT2006-90166.

This paper reports on the validation of the assumption of quasi steady behaviour of pulsating cooling injection in the near hole flow region. The respective experimental data are taken in a flat plate wind tunnel at ETH Zürich. The facility simulates the film cooling row flow field on the pressure side of a turbine blade. Engine representative non-dimensionals are achieved, providing a faithful model at larger scale. Heating the free stream air and strongly cooling the coolant gives the required density ratio between coolant and free-stream. The coolant is injected with different frequency and amplitude. The three dimensional velocities are recorded using non-intrusive PIV, seeding is provided for both air streams. Two different cylindrical hole geometries are studied, with different angles. Blowing ratio is varied over a range to simulate pressure side film cooling. The general flow field, the jet trajectory and the streamwise circulation are utilized in the validation of the quasi steady assumption.

Topics: Cooling , Modeling
Commentary by Dr. Valentin Fuster
2006;():139-148. doi:10.1115/GT2006-90168.

Exit surveys detailing total pressure loss, turning angle, and secondary velocities have been acquired for a fully loaded vane profile in a large scale low speed cascade facility. Exit surveys have been taken over a four-to-one range in Reynolds numbers based on exit conditions and for both a low turbulence condition and a high turbulence condition. The high turbulence condition was generated using a mock aero-derivative combustor. Exit loss, angle, and secondary velocity measurements were acquired in the facility using a five-hole cone probe at two stations representing axial chord spacings of 0.25 and 0.50. Substantial differences in the level of losses, distribution of losses, and secondary flow vectors are seen with the different turbulence conditions and at the different Reynolds numbers. The higher turbulence condition produces a significantly broader wake than the low turbulence case and shows a measurable total pressure loss in the region outside the wakes. Generally, total pressure losses are about 0.02 greater for the high turbulence case compared with the low turbulence case primarily due to the state of the suction surface boundary layers. Losses decrease moderately with increasing Reynolds number. Cascade inlet velocity distributions have been previously documented in an endwall heat transfer study of this same geometry. These exit survey measurements support our understanding of the endwall heat transfer distributions, the secondary flows in the passage, and the origin of losses.

Commentary by Dr. Valentin Fuster
2006;():149-159. doi:10.1115/GT2006-90170.

Full surface pressure distributions over the endwall and pin in a staggered pin fin array have been acquired over a ten to one range in Reynolds numbers. These pressure distributions allow us to visualize the strong inertial pressure gradients that are responsible for driving secondary flows in pin fin passages. These strong pressure gradients include endwall regions near the pin stagnation region and near the pin at 90° from the stagnation region. Pressure distributions have been acquired on pin and endwall surfaces at eight consecutive rows using conventional static pressure measurement techniques. Pressures have been taken at 380 locations per row and, assuming symmetry, provide a well resolved visualization of surface pressure. Generally, surface and pin pressure distributions vary significantly from row to row in the entrance of the array at a given Reynolds number but stay relatively consistent after row four. Dimensionless pressure distributions are quite similar for row one for all Reynolds numbers but vary significantly at a given row downstream with Reynolds number. These data are expected to enhance our understanding of pin array fluid dynamics and to compliment full surface heat transfer data presented in a future paper.

Commentary by Dr. Valentin Fuster
2006;():161-172. doi:10.1115/GT2006-90173.

Within a European research project the tip end wall region of LP turbine guide vanes with leakage ejection was investigated at DLR in Göttingen. For this purpose a new cascade wind tunnel with three large profiles in the test section and a contoured end wall was designed and built up, representing 50% height of a real low pressure turbine (LPT) stator and simulating the casing flow field of shrouded vanes. The effect of tip leakage flow was simulated by blowing air through a small leakage gap in the end wall just upstream of the vane leading edges. Engine relevant turbulence intensities were adjusted by an active turbulence generator mounted in the test section inlet plane. The experiments were performed with tangential and perpendicular leakage ejection and varying leakage mass flow rates up to 2%. Aerodynamic and thermodynamic measurement techniques were employed. Pressure distribution measurements provided information about the end wall and vane surface pressure field and its variation with leakage flow. Additionally streamline pattern (local shear stress directions) on the walls were detected by oil flow visualization. Downstream traverses with 5-hole pyramid type probes allow a survey of the secondary flow behavior in the cascade exit plane. The flow field in the near end wall area downstream of the leakage gap and around the vane leading edges was investigated using a 2D Particle Image Velocimetry (PIV) system. In order to determine end wall heat transfer distributions, the wall temperatures were measured by an infra-red camera system, while heat fluxes at the surfaces were generated with electric operating heating foils. It turned out from the experiments that distinct changes in the secondary flow behavior and end wall heat transfer mainly occur when the leakage mass flow rate is increased from 1% to 2%. Leakage ejection perpendicular to the main flow direction amplifies the secondary flow, in particular the horse-shoe vortex, whereas tangential leakage ejection causes a significant reduction of this vortex system. For high leakage mass flow rates the boundary layer flow at the end wall is strongly affected and seems to be highly turbulent, resulting in entirely different heat transfer distributions.

Commentary by Dr. Valentin Fuster
2006;():173-182. doi:10.1115/GT2006-90177.

The objective of this work is to compare the predicted flow field and the endwall heat transfer of a baseline nozzle guide vane configuration with a combustion chamber variant, a heat shield variant without and with additional cooling air, and a cavity variant without and with additional cooling air. The comparison is carried out numerically using the commercial 3D Navier-Stokes software package Fluent [1]. For the turbulence modeling the v2 -f model by Durbin [2] been used. The detailed comparison of the flow field and the endwall heat transfer shows major differences between the baseline and heat shield configuration. The heat shield in front of the airfoil of the nozzle guide vane cascade influences the secondary flow field and the endwall heat transfer pattern strongly. The additional cooling air, released under the heat shield also has a distinctive influence. The cavity between the combustion chamber and the nozzle guide vane affects the secondary flow field and the endwall heat transfer pattern. Here the influence of the additional cavity cooling air is more decisive.

Commentary by Dr. Valentin Fuster
2006;():183-187. doi:10.1115/GT2006-90196.

Investigation of local heat transfer characteristics near a row of film cooling holes in the inner side of a simulated turbine blade midchord region with impingement has been carried out experimentally. The research about heat transfer characteristics is focused on three diameter of film cooling hole area located upstream and downstream a row of film cooling holes, which angle is at a 90 degrees. The internal impingement air is provided by a single line of equally spaced jets. The film cooling air extracts through a line of holes on the impinging target plate. The projection of the jets on the target plate is always on the center line between two film holes. The spacing of the jet holes is twice that of the film cooling holes. The effect of the streamwise arrangement of the impingement nozzles relative to the position of the film cooling holes and impinging distance on the heat transfer characteristics have been mainly investigated. The experiment is conducted under the flow condition of Reynolds number 10000∼30000, crossflow-to-jet mass flux ratio based on each channel/jet hole section area 0.1 and film outflow-to-crossflow mass flux ratio based on film cooling hole/channel section 12∼20. In the range of experimental parameter, the experimental results indicate that there is optimal ratio of the impinging distance to film hole diameter, on which the heat transfer characteristics is best. Similarly for the area upstream film cooling hole, there is the optimal ratio of distance of the impingement nozzles relative to the position of the film cooling holes to film hole diameter. As impinging holes are away from film cooling holes in the streamwise direction of crossflow, the effect of impingement on local heat transfer near film cooling holes is weakened, but film cooling extraction effect stand out. The place closer to the hole will have stronger heat transfer whether upstream the film cooling holes or downstream the holes. Based on this, the effects of position of the jets relative to the film cooling holes on the heat transfer characteristics have been obtained qualitatively and quantitatively. It can be the important reference for accurately designing gas turbine blade.

Commentary by Dr. Valentin Fuster
2006;():189-194. doi:10.1115/GT2006-90197.

Experimental investigations of local impingement heat transfer characteristics near a row of film cooling holes in a simulated internal midchord region of gas turbine blade have been carried out. The research of heat transfer characteristics is focused on three film cooling hole diameter area located upstream and downstream a row of film holes. There is a line of equally spaced film cooling holes whose angles are 30 or 90 degrees. When there is no impingement, the investigation about the effect of the film cooling bleed has been carried out under different cross flow Reynolds Numbers and film outflow-to-crossflow mass flux ratios based on each film cooling hole/channel section area. The results indicate that the local heat transfer near the film cooling holes is enhanced with the increase of the crossflow Reynolds Numbers and film outflow-to-crossflow mass flux ratios based on each film cooling hole/channel-section area. The local heat transfer characteristic downstream film cooling holes is better than that upstream film cooling holes. The average Nusselt number of one time diameter area downstream the row of film holes is generally 40% more than that upstream the row of film cooling holes. The place closer to the hole will have stronger heat transfer whether upstream film cooling holes or downstream film cooling holes. When there is impingement, the impinging air is provided by a single line of equally spaced jets. The spacing of the jet holes is twice that of the film cooling holes with staggered arrangements. The local heat transfer near the row of film cooling holes has been studied experimentally through changing flow parameters, such as impinging Reynolds Numbers and mass flux ratios of crossflow-to-jet based on each channel/jet hole section area etc. A great number of experimental data has been obtained. Based on this, the effects of the flow parameters on the heat transfer characteristics have been obtained qualitatively and quantitatively. It can be the important reference for accurately designing gas turbine blade.

Commentary by Dr. Valentin Fuster
2006;():195-203. doi:10.1115/GT2006-90225.

Film cooling adiabatic effectiveness for axial and compound angle holes on the suction side of a simulated turbine vane was investigated to determine the relative performance of these configurations. The effect of the surface curvature was also evaluated by comparing to previous curvature studies and flat plate film cooling results. Experiments were conducted for varying coolant density ratio, mainstream turbulence levels, and hole spacing. Results from these measurements showed that for mild curvature, 2r/d ≈ 160, flat plate results are sufficient to predict the cooling effectiveness. Furthermore, the compound angle injection improves adiabatic effectiveness for higher blowing ratios, similar to previous studies using flat plate facilities.

Commentary by Dr. Valentin Fuster
2006;():205-213. doi:10.1115/GT2006-90226.

Adiabatic film cooling effectiveness of axial holes embedded within a transverse trench on the suction side of a turbine vane was investigated. High resolution two dimensional data obtained from IR thermography and corrected for local conduction provided spatial adiabatic effectiveness data. Flow parameters of blowing ratio, density ratio, and turbulence intensity were independently varied. In addition to a baseline geometry, nine trench configurations were tested, all with a depth of 1/2 hole diameter, with varying widths, and with perpendicular and inclined trench walls. A perpendicular trench wall at the very downstream edge of the coolant hole was found to be the key trench characteristic that yielded much improved adiabatic effectiveness performance. This configuration increased adiabatic effectiveness up to 100% near the hole and 40% downstream. All other trench configurations had little effect on the adiabatic effectiveness. Thermal field measurements confirmed that the improved adiabatic effectiveness that occurred for a narrow trench with perpendicular walls was due a lateral spreading of the coolant and reduced coolant jet separation. The cooling levels exhibited by these particular geometries are comparable to shaped holes, but much easier and cheaper to manufacture.

Commentary by Dr. Valentin Fuster
2006;():215-229. doi:10.1115/GT2006-90229.

This paper concerns itself with investigating the effect of rotation on flow and heat transfer in a 45° ribbed square duct. Large-Eddy Simulations (LES) are used to investigate why rotation does not have any effect on heat transfer augmentation unlike 90 degree ribs, in which considerable changes are observed in augmentation at the trailing and leading walls of the duct. It is found that unlike 90 degree ribbed ducts, in which the heat transfer augmentation is strongly dependent on streamwise momentum, spanwise momentum dominates heat transfer in skewed ribs. Since Coriolis forces under orthogonal rotation about the z-axis do not directly contribute to spanwise momentum, they do not have as much of an effect on heat transfer at the ribbed walls at the trailing and leading sides. However, because of the augmentation of turbulence at the trailing side, the vortices which are produced in the separated shear layer of the rib and which move from the inside to the outside of the duct, break down and diffuse before they can impinge on the outer wall. Turbulence attenuation at the leading wall has the opposite effect which allows the vortices to maintain their coherence and impinge on the outer wall. This effect taken together with the streamwise flow being pushed to the leading side, produces an extended region of high heat transfer at the outer wall near the leading side. This is countered by lower heat transfer at the trailing side of the outer wall. Hence, although local variations are present due to rotation, the overall augmentation remains the same.

Commentary by Dr. Valentin Fuster
2006;():231-240. doi:10.1115/GT2006-90234.

Numerical investigations on the film cooling of an inlet guide vane are performed with realistic geometry. The vane model comprises one vane passage, 131 shower-head cooling holes in 6 staggered rows around the vane leading edge, and a coolant supply plenum. A fully implicit coupled 3D N-S solver based on finite-volume method and incorporated with unstructured mixed grid, standard k–ε turbulence model and scalable wall function is employed to obtain the numerical solution. Two film cooling configurations, named original design and modified design, are presented. The original design and no cooling case are simulated to obtain flow mechanism and heat transfer characteristics of the leading edge film cooling. In addition, the effects of the meridional endwall contours on the leading edge film cooling are considered. The film cooling characteristics and interactions between jets and mainstream around the leading edge, especially near the stagnation line, are analyzed in detail. To provide better coolant coverage on the leading edge, the cooling configuration is modified by redistributing the position and direction of some rows of holes based upon the analysis and understanding of the 3D prediction for the original design. The modified design is verified under three blowing ratios and compared with the original design.

Commentary by Dr. Valentin Fuster
2006;():241-247. doi:10.1115/GT2006-90250.

An experimental study was carried out to comprehend the passage configuration for better the cooling effectiveness on the trailing edge of high-pressure turbine blade. Thermochromic liquid crystal technology was employed to measure the endwall temperature of the trapezoidal compound passage and the resistance coefficients of the internal ducts. It can be concluded that the configuration with suitable ejection holes in the divider wall can get higher heat transfer coefficient of the compound passage than that of the tradition configuration with a straight divider wall which form a sharp turn. This is because the spanwise impinging jets flowing through the orifices distributed at the radial orientation in the divider wall between the first passage and the second one, can improve the heat transfer of a certain low heat transfer region. Better effectiveness of heat transfer could be obtained when the straight divider wall was designed into zigzag divider wall.

Commentary by Dr. Valentin Fuster
2006;():249-257. doi:10.1115/GT2006-90269.

Effective cooling of combustor liners is an important and integrated part of jet engine combustor design, as the liner temperature directly impacts the durability of the combustor. Cooling air jet, usually around 1000–1200 °F, is inserted into slots at the onset of each liner panel. These slots and feedholes are together called nuggets. The discrete air jets are expected to coalesce and form a cooling film with uniform velocity at the exit of the slots. This paper presents results of CFD analysis using realizable k–ε turbulence model in a commercial CFD software FLUENT®. For accurate analysis of different nugget feedhole configurations, an appropriately selected turbulent model is crucial for flow and heat transfer analysis. Special care is given to the coefficient of destruction term related to the turbulent decay exponent of the dissipation rate in transport equation. Validation is achieved by comparison of predicted and measured liner cooling film effectiveness from experimental tests conducted under various blowing ratios. Mass transfer analogy method is used to measure cooling film effectiveness, represented by mass fraction of CO2 along combustor liner. Reasonably good agreement is obtained between the CFD analysis results and the experimental data.

Commentary by Dr. Valentin Fuster
2006;():259-269. doi:10.1115/GT2006-90272.

The effects of the coolant jet pulsing frequency (PF), duty cycle (DC), and hole shape geometry on heat transfer coefficient and film effectiveness were investigated with a film hole located on a semicircular leading edge test model with an afterbody. Cylindrical and diffusion-shaped holes located at 21.5° from the stagnation line were investigated. An infrared thermography technique with a single transient test was used to determine both the heat transfer coefficient and film effectiveness. Spanwise averaged heat transfer coefficient and film effectiveness were computed from the local values for all test conditions under the same Reynolds number (Re) of 60,000 and density ratio (DR) of 1.11. A dimensionless Frossling number (Fr) was used to represent the heat transfer coefficient. The effects of duty cycles of 50%, 75%, and 100% (continuous coolant) on film effectiveness and heat transfer coefficient were investigated at coolant jet pulsing frequencies of 5 Hertz (Hz) and 10 Hertz. The duty cycle and pulsing frequency were controlled by the opening and closing time settings of two synchronized pulsed valves. The blowing parameters investigated included continuous coolant at the blowing ratios (M) of 0.75, 1.00, 1.50 and 2.00. The subsequent pulsed cases for a combination of pulsing frequency and duty cycle were varied from the corresponding continuous case without changing the coolant flow rate (or blowing ratio) setting for a total of 40 cases for the shaped and cylindrical film holes. The shaped hole provides higher local film effectiveness values than the classical cylindrical hole when coolant flow is steady at M = 1.00. The higher local film effectiveness for the shaped hole was also observed for pulsed cases at M = 1.50 (Meff = 1.25) and M = 2.00 (Meff = 1.07) due to wider film spreading or coverage. The pulsed coolant cases provide higher spanwise averaged film effectiveness than the continuous coolant at M = 1.50 for both hole geometries. In contrast to the film effectiveness, the spanwise averaged Frossling numbers of pulsed coolant are lower compared to the continuous coolant for both hole shapes at the same blowing ratio. Combining the effects of heat transfer coefficient and film effectiveness, one can compute a relative heat load ratio to evaluate the performance of the film cooling. The pulsed coolant cases in general perform better than continuous coolant. The shaped hole geometry provides better film cooling performance than the cylindrical hole geometry for all blowing ratios including the continuous and the pulsed coolant cases studied.

Commentary by Dr. Valentin Fuster
2006;():271-283. doi:10.1115/GT2006-90276.

Historically the design of gas turbine engines have not considered the interaction between the combustor and turbine stages. High pressure turbine vane stages have been designed assuming inlet conditions consistent with a standard turbulent boundary layer profile. However, combustor exit flow entering the vane is known to be highly non-uniform in both the primary and secondary flow regimes. In order to develop higher performance, more efficient, longer life stages, turbine design must take into account combustor exit non-uniformities. The Turbine Research Facility (TRF) at Wright-Patterson Air Force Base has installed a non-reactive full scale annular combustor simulator or more accurately a turbine inlet profile generator to study combustor-vane interaction. Several benchmark tests have been performed on the profile generator consisting of a Taguchi type matrix wherein nine independent variables were adjusted. Supplementing the experimental research at the TRF, a steady state, unstructured, fully three-dimensional CFD analysis was performed. This paper will make comparisons between the CFD and experimental profiles generated by the simulator. Furthermore, the computational study will help to give an understanding of the aerodynamic and aerothermal environment within the generator that experimental instrumentation alone cannot.

Commentary by Dr. Valentin Fuster
2006;():285-295. doi:10.1115/GT2006-90277.

The goal of this work was to investigate the effects of different profiles representative of those exiting aero-engine combustors on high pressure turbine vane aerodynamics and heat transfer. The various profiles were produced using the non-reacting, inlet profile generator in the Turbine Research Facility (TRF) located at the Air Force Research Laboratory (AFRL). This paper reports how the pressure loading and heat transfer along the vane surface is affected by different turbine inlet pressure and temperature profiles at several different span locations. The results indicate that the different inlet total pressure profiles affected the aerodynamic loading by as much as 10%. The results also reveal that the combination of different total pressure and total temperature profiles significantly affected the vane heat transfer for a baseline test with relatively uniform inlet total pressure and total temperature profiles. Near the ID endwall, the baseline heat transfer was reduced 30 to 40% over the majority of the vane surface. Near the OD endwall, it was found that certain inlet profiles could increase the baseline heat transfer by 20 to 30%, while other profiles resulted in a decrease of the baseline heat transfer by 30 to 40%.

Commentary by Dr. Valentin Fuster
2006;():297-305. doi:10.1115/GT2006-90301.

A model for the diffusion of turbulent kinetic energy (TKE) for high free stream turbulence (FST) boundary layers is implemented in commercially available FLUENT CFD code to predict the effects of high free stream turbulence on heat and momentum transport in a flat plate boundary layer using Launder and Spalding’s standard k-ε model. The computational results are compared with experimental data sets. When experimental and/or standard initial profiles and standard k-ε model were used for calculations of Stanton number and skin friction coefficient under high FST intensities the results were close to the experimental data (within 2%). However. TKE profiles had large deviations (within approximately 40 %) compared to the data for both moderately high FST (Tui = 6.53%) and very high FST (Tui = 25.7%) intensities. Since TKE values are used in calculations of skin friction coefficients and Stanton numbers through calculation of turbulent viscosity from k and ε, getting a correct result for these quantities from the wrong calculations of TKE seems contradictory. In an earlier study it was concluded that the TKE calculations were low compared to the data because k-ε models do not model the diffusion of high FST correctly. To correct this discrepancy, a new model for TKE diffusion was developed and used it in TEXSTAN code. The objective of the current study is to generalize this model and use it in more complicated geometries by applying to FLUENT code. Therefore at this first phase of this study, this diffusion model was implemented in the Launder and Spalding k-ε model contained in FLUENT code using User Defined Functions (UDF) by modifying the turbulent kinetic energy transport equation. The constant Cμ which exists in the turbulent viscosity equation was also modified using experimental data. This model considerably increased the TKE values for both moderately high and very high FST intensities showing that it functions as it was intended. While TKE perfectly matched with the experimental data sets (within 1–2%) for moderately high initial FST intensity, it still did not yield very good results for very high initial FST intensity. Under very high FST intensity TKE does not match data very well near the wall. The Stanton number and skin friction coefficient increased (about 30%) as expected since the new diffusion model increases TKE levels ner the wall. At this point it should be mentioned that standard k-ε model is a high Reynold number turbulence model which use wall functions near the wall. In earlier studies the new diffusion model was applied to a low Reynold number model in TEXSTAN code. In the continuing studies high Reynolds number k-ε model in FLUENT will be modified to create a low Reynolds number model via UDFs in order to get better predictions of the Stanton number and skin friction coefficients by use of damping function fμ and adjusting the value of turbulent Prandtl number. This study reports on progress in overall goal of implementing a new TKE diffusion model in FLUENT code.

Commentary by Dr. Valentin Fuster
2006;():307-313. doi:10.1115/GT2006-90321.

To enhance the film cooling performance in the vicinity of the turbine blade leading edge, the flow characteristics of the film-cooled turbine blade have been investigated using a cylindrical body model. The inclination of the cooling holes is along the radius of the cylindrical wall and 20 deg relative to the spanwise direction. Mainstream Reynolds number based on the cylinder diameter was 1.01×105 and 0.69×105 , and the mainstream turbulence intensities were about 0.2% in both Reynolds numbers. CO2 was used as coolant to simulate the effect of density ratio of coolant-to-mainstream. Furthermore, the effect of coolant flow rates was studied for various blowing ratios of 0.4, 0.7, 1.1, and 1.4, respectively. In experiment, spatially-resolved temperature distributions along the cylindrical body surface were visualized using infrared thermography (IRT) in conjunction with thermocouples, digital image processing, and in situ calibration procedures. This comparison shows the results generated to be reasonable and physically meaningful. The film cooling effectiveness of current measurement (0.29 mm × 0.33 min per pixel) presents high spatial and temperature resolutions compared to other studies. Results show that the blowing ratio has a strong effect on film cooling effectiveness and the coolant trajectory is sensitive to the blowing ratio. The local spanwise-averaged effectiveness can be improved by locating the first-row holes near the second-row holes.

Commentary by Dr. Valentin Fuster
2006;():315-324. doi:10.1115/GT2006-90322.

Numerical simulations were performed to predict the effect of cavity purge flow on the rotating blade platform in a 1-1/2 turbine stage using a Reynolds stress turbulence model together with a non-equilibrium wall function. Simulations were carried out with a sliding mesh for the rotor under three rotating speeds (2000, 2550 and 3000 rpm) and three purge-to-mainstream mass flow ratios (0.5%, 1% and 1.5%) to investigate the effects of rotating speed and coolant purging rate on the rotating blade platform film cooling. The adiabatic film cooling effectiveness was evaluated using the adiabatic wall temperatures with and without coolant purging to examine the true effect of coolant protection. The film cooling effectiveness increases with increasing coolant purging flow ratio from 0.5% to 1.5% of mainstream. Higher rotating speed also enhances film cooling effectiveness for the range of rotating speed considered. The predicted laterally averaged adiabatic film cooling effectiveness is in good agreement with the corresponding experiment data except for the platform leading edge region. However, the detailed effectiveness distribution on the platform is not well predicted by this study. In addition, the detailed instantaneous film cooling effectiveness and the associated heat transfer coefficients for four different time phases are also reported.

Commentary by Dr. Valentin Fuster
2006;():325-340. doi:10.1115/GT2006-90352.

In certain regions of turbine aerofoils, cooling system designers need to cool the blades with convection systems that provide high heat transfer coefficients. The present research has investigated a circular cooling passage with tangential injection suitable for a blade leading edge. The heat transfer coefficients are measured using the conventional transient heat transfer, liquid crystal technique. The results are compared to the data from steady state experiments performed by Hedlund et al. [1]. The cooling system performance is compared in detail to average data from earlier tangential injection experiments and to local heat transfer coefficient expected from a normal impingement system. The vortex flow field was also studied by numerical prediction and near-wall velocity measurements. The investigation of the flow structure has led to understanding of flow mechanisms responsible for the high heat transfer coefficient. The vortex flow field was also investigated using computational fluid dynamics and with hot wire anemometry. The latter near wall measurements were combined with the law of the wall and Colburn analogy to validate the flow and heat transfer measurements.

Commentary by Dr. Valentin Fuster
2006;():341-351. doi:10.1115/GT2006-90355.

Thermodynamic and aerodynamic measurements at and near the endwall of turbine vanes were carried out in a linear cascade with a transonic flow field. The investigations were performed in the Windtunnel for Straight Cascades at DLR Göttingen at representative dimensionless engine conditions of Mach and Reynolds number, Ma2is = 1.0 and Re2 = 850 000 respectively. The endwall film cooling configuration consisted of a slot in front of the vanes, film cooling holes inside the vane passages and a groove simulating the slit between two adjacent vane platforms, but there was no coolant leakage from the groove. Laser-Two-Focus velocimetry (L2F) was used to determine local velocities in the vicinity of the endwall. At a much larger number of locations compared to the velocity measurements the L2F-device was utilized as a seeding particle counter which enabled the determination of local coolant concentration. With these concentration measurements the migration of coolant from the different origins could be traced through the vane passage. By extrapolating the measured concentration values to the endwall adiabatic film cooling effectiveness could be obtained. The measurements at the present slot configuration were compared with previous ones [1, 2] where the slot position was closer to the vane entrance. Whereas the coolant ejection at the previous slot position produced a much more intense horse shoe vortex than without coolant, the new slot position causes no increase of secondary flow. This result proves the previous statement that positioning a coolant opening flush near the saddle point of the upstream endwall boundary layer stagnation region should be avoided. The new slot position improved film cooling effectiveness compared to the previous ones even with half the amount of coolant. By investigating the migration of film coolant from the holes inside the vane passage, ineffective holes could be identified and suggestions for improving the film coolant configuration could be given. At one location adiabatic film cooling effectiveness from these aerodynamic measurements could be compared with a thermodynamic measurement using infrared imaging.

Topics: Coolants , Turbines
Commentary by Dr. Valentin Fuster
2006;():353-361. doi:10.1115/GT2006-90367.

To optimize turbine blade showerhead film cooling, detailed film effectiveness was measured for four different showerhead geometries in a warm cascade simulating realistic engine operation conditions. Local film effectiveness distributions were obtained on both the pressure and suction surfaces of blade models using the pressure sensitive paint (PSP) technique. The four different geometries that have been investigated include: baseline geometry with a three-row showerhead; reduced injection angle geometry; a two-row geometry and increased diameter geometry. Nitrogen gas was used to simulate cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness by the mass transfer analogy could be obtained. Cooling mass flow was controlled to have a coherent comparison between the different geometries. Local film effectiveness distributions were measured for each of the cooling mass flows. Then the distributions were spanwise averaged for comparison. Reducing the injection angle or increasing the hole diameter, the film effectiveness improved slightly for a fixed total coolant flow. The two-row injection resulted in poor film effectiveness distribution possibly due to jet lift-off as it had higher momentum compared to the three-row injection.

Commentary by Dr. Valentin Fuster
2006;():363-373. doi:10.1115/GT2006-90368.

Rib turbulators are commonly used to enhance the heat transfer within internal cooling passages of advanced gas turbine blades. Many factors affect the thermal performance of a cooling channel with ribs. This study experimentally investigates the effect of rib spacing on the heat transfer enhancement, pressure penalty, and thus the overall thermal performance in both rotating and non-rotating rectangular, cooling channels. In the 1:2 rectangular channels, 45° angled ribs are placed on the leading and trailing surfaces. The pitch of the ribs varies, so rib pitch-to-height (P/e) ratios of 10, 7.5, 5, and 3 are considered. Square ribs with a 1.59 mm × 1.59 mm cross-section are used for all spacings, so the height-to-hydraulic diameter (e/Dh ) ratio remains constant at 0.094. With a constant rotational speed of 550 rpm and the Reynolds number ranging from 5000 to 40000, the rotation number in turn varies from 0.2 to 0.02. Because the skewed turbulators induce secondary flow along the length of the rib, the very close rib spacing of P/e = 3, has the best thermal performance in both rotating and non-rotating channels. This close spacing yields the greatest heat transfer enhancement, while the P/e = 5 spacing has the greatest pressure penalty. In addition, the effect of rotation is more pronounced in the channel with the rib spacing of 3. As more ribs are added, the channel is approaching a smooth channel, and the strength of the rotation induced vortices increases.

Commentary by Dr. Valentin Fuster
2006;():375-382. doi:10.1115/GT2006-90370.

Detailed film cooling measurements are presented on a turbine blade leading edge model with three rows of showerhead holes. Experiments are run at a mainstream Reynolds number of 19,500 based on cylindrical leading edge diameter. One row of holes is located on the stagnation line and the other two rows are located at ±15° on either side of the stagnation line. The three rows have compound angle holes angled 90° in the flow direction, 30° along the spanwise direction, and the two holes on either side of the stagnation row have and additional angle of 0°, 30°, and 45° in the transverse direction. The effect of hole shaping of the 30° and 45° holes is also considered. Detailed heat transfer coefficient and film effectiveness measurements are obtained using a transient infrared thermography technique. The results are compared to determine the advantages of shaping the compound angle for rows of holes off stagnation row. Results show that, the additional compound angle in the transverse direction for the two rows adjacent to the stagnation row provide significantly higher film effectiveness than the typical leading edge holes with only two angles. Results also show that, the shaping of showerhead holes provides higher film effectiveness than just adding an additional compound angle in the transverse direction and significantly higher effectiveness than the baseline typical leading edge geometry. Heat transfer coefficients are higher as the spanwise angle for this study is larger than typical leading edge geometries with an angle of 30° compared to 20° for other studies.

Topics: Cooling
Commentary by Dr. Valentin Fuster
2006;():383-394. doi:10.1115/GT2006-90375.

A five blade, linear cascade is used to experimentally investigate turbine blade platform cooling. A 30° inclined slot upstream of the blades is used to model the seal between the stator and rotor, and 12 discrete film holes are located on the downstream half of the platform for additional cooling. The film cooling effectiveness is measured on the platform using pressure sensitive paint (PSP). The mainstream Reynolds number is 3.1*105 based on the inlet velocity and the chord length of the scaled high pressure turbine blade. The upstream slot covers 1.5 passages with the coolant exiting the slot at the leading edge of the rotor blades. The length-to-width ratio (ls /w) of the slot is 5.7, and the slot flowrate varies from 0.50% to 2.0% of the mainstream flow. The discrete film cooling holes also have an inclination of 30°, so the length-to-diameter (lf /d) ratio of each hole is 10. The blowing ratio of the coolant through the holes varies from 0.5 to 2.0, based on the mainstream exit velocity. Using PSP it is clear that the film cooling effectiveness on the blade platform is strongly influenced by the platform secondary flow through the passage. Increasing the slot injection rate weakens the secondary flow and provides more uniform film coverage. Increasing the freestream turbulence level was shown to increase film cooling effectiveness on the endwall, as the increased turbulence also weakens the passage vortex. However, downstream, near the discrete film cooling holes, the increased turbulence decreases the film cooling effectiveness (as reported for flat plate film cooling studies). Finally, combining upstream slot flow with downstream discrete film holes should be done cautiously to ensure coolant is not wasted by overcooling regions on the platform.

Commentary by Dr. Valentin Fuster
2006;():395-406. doi:10.1115/GT2006-90377.

Three different designs of a transpiration cooled multilayer plate — plane, convex and concave — are analysed numerically by application of a 3-D conjugate fluid flow and heat transfer solver. The geometrical setup and the fluid flow conditions are derived from modern gas turbine components. The conjugate analysis of these designs focus on the influence of the surface curvature, the cooling film development on the plate surface, the fluid structure in the cooling channels and on the cooling efficiency of the plate. Moreover, to predict the effective thermal properties and the permeability of these multilayer plates, a multiscale approach based on the homogenization technique is employed. This method allows the calculation of effective equivalent properties either for each layer or for the multilayer of superalloy, bondcoat and thermal barrier coating (TBC). Permeabilities of the different designs are presented in detail for the TBC layer. The influence of the plate curvature and the blowing ratio on the effective orthotropic thermal conductivities is finally outlined.

Commentary by Dr. Valentin Fuster
2006;():407-418. doi:10.1115/GT2006-90379.

With the increase in usage of gas turbines for power generation and given that natural gas resources continue to be depleted, it has become increasingly important to search for alternate fuels. One source of alternate fuels is coal derived synthetic fuels. Coal derived fuels, however, contain traces of ash and other contaminants that can deposit on vane and turbine surfaces affecting their heat transfer through reduced film-cooling. The endwall of a first stage vane is one such region that can be susceptible to depositions from these contaminants. This study uses a large-scale turbine vane cascade in which the following effects on film-cooling adiabatic effectiveness were investigated in the endwall region: the effect of near-hole deposition, the effect of partial film-cooling hole blockage, and the effect of spallation of a thermal barrier coating. The results indicated that deposits near the hole exit can sometimes improve the cooling effectiveness at the leading edge, but with increased deposition heights the cooling deteriorates. Partial hole blockage studies revealed that the cooling effectiveness deteriorates with increases in the number of blocked holes. Spallation studies showed that for a spalled endwall surface downstream of the leading edge cooling row, cooling effectiveness worsened with an increase in blowing ratio.

Commentary by Dr. Valentin Fuster
2006;():419-432. doi:10.1115/GT2006-90390.

The design of a three-dimensional non-axisymmetric end-wall is carried out using three-dimensional numerical simulations. The computations have been conducted both for the flat and contoured end-walls. The performance of the end-wall is evaluated by comparing the heat transfer and total pressure loss reduction. The contouring is done in such a way to have convex curvature in the pressure side and concave surface in the suction side. The convex surface increases the velocity by reducing the local static pressure while concave surface decreases the velocity by increasing the local pressure. The profiling of the end-wall is done by combining two curves, one that varies in the streamwise direction while the other varies in the pitchwise direction. Several contoured end-walls are created by varying the streamwise variation keeping the pitchwise curve constant. The flow near the contoured end-wall is seen to be significantly different than that of flat end-wall. The contoured end-wall is found to reduce the secondary flow by decreasing radial pressure gradient. The total pressure loss is also lower and the average heat transfer reduces by about 8% compared to the flat end-wall. Local reductions in heat transfer are significant (factor of 3). This study demonstrates the potential of three-dimensional end-wall contouring for reducing the thermal loading on the end wall.

Commentary by Dr. Valentin Fuster
2006;():433-445. doi:10.1115/GT2006-90391.

Heat transfer measurements are reported for a rotating 4:1 aspect ratio (AR) coolant passage with ribs skewed 45 degree to the flow. The study covers Reynolds number (Re) in the range of 10,000–70,000, rotation number (Ro) in the range of 0–0.6, and density ratios (DR) between 0.1–0.2. These measurements are done in a rotating heat transfer rig utilizing segmented copper pieces that are individually heated, and thermocouples with slip rings providing the interface between the stationary and rotating frames. The results are compared with the published data obtained in a square channel with similar dimensionless rib-geometry parameters, and with the results obtained for a 4:1 AR smooth channel. As in a 1:1 AR channel, rotation enhances the heat transfer on the destabilized walls (inlet-trailing wall and outlet-leading wall), and decreases the heat transfer ratio on the stabilized walls (inlet-leading wall and outlet-trailing wall). However, the rotation-induced enhancement/degradation for the 4:1 rectangular channel is much weaker than that in the square ribbed channel, especially in the inlet (the first passage). The results on the inlet-leading wall are in contrast to that in the smooth channel with the same AR, where rotation causes heat transfer to increase along the inlet-leading wall at lower Reynolds number (Re = 10,000 and 20,000). Higher DR is observed to enhance the heat transfer on both ribbed walls in the inlet (the first passage) and the outlet (the second passage), but the DR effects are considerably weaker than those in a ribbed square channel. Measurements have also been parameterized with respect to the buoyancy parameter and results show the same general trends as those with respect to the rotation number. In addition, pressure drop measurements have been made and the thermal performance factor results are presented.

Commentary by Dr. Valentin Fuster
2006;():447-458. doi:10.1115/GT2006-90401.

Improving the performance and durability of gas turbine aircraft engines depends highly on achieving a better understanding of the flow interactions between the combustor and turbine sections. The flow exiting the combustor is very complex and it is characterized primarily by elevated turbulence and large variations in temperature and pressure. The heat transfer and aerodynamic losses that occur in the turbine passages are driven primarily by these spatial variations. To better understand these effects, the goal of this work is to benchmark an adjustable turbine inlet profile generator for the Turbine Research Facility (TRF) at the Air Force Research Laboratory (AFRL). The research objective was to experimentally evaluate the performance of the non-reacting simulator that was designed to provide representative combustor exit profiles to the inlet of the TRF turbine test section. This paper discusses the verification testing that was completed to benchmark the performance of the generator. Results are presented in the form of temperature and pressure profiles as well as turbulence intensity and length scale. This study shows how one combustor geometry can produce significantly different flow and thermal field conditions entering the turbine. Engine designers should place emphasis on obtaining accurate knowledge of the flow distribution within the combustion chamber as this can result in significantly different inlet profiles to the turbine that can change local aerodynamics and heat transfer within the turbine.

Commentary by Dr. Valentin Fuster
2006;():459-468. doi:10.1115/GT2006-90405.

Gas turbine cooling has steadily acquired major importance whenever engine performances have to be improved. Among various cooling techniques, film cooling is probably one of the most diffused systems for protecting metal surfaces against hot gases in turbine stages and combustor liners. Most recent developments in hole manufacturing allow to perform a wide array of micro-holes, currently referred to as effusion cooling. This paper presents the validation of a simplified 2D conjugate approach through comparison with the experimental results of effectiveness for an effusion plate, performed during the first year of the European Specific Targeted REsearch Project AITEB-2 (Aerothermal Investigation of Turbine Endwalls and Blades). A preliminary test is performed with the steady-state technique, using TLC (Thermochromic Liquid Crystal) wide-band formulations. Results are obtained in terms of local distributions of adiabatic effectiveness. Average values are compared with calculations to validate the numerical code. Then, Design Of Experiment (DOE) approach is used to perform several conjugate tests (about 180), so as to derive the behavior of different effusion plates in terms of overall effectiveness and mass flow rate. Data are analyzed in detail and a correlative approach for the overall effectiveness is proposed.

Commentary by Dr. Valentin Fuster
2006;():469-482. doi:10.1115/GT2006-90425.

High resolution Nusselt number (Nu) distributions were measured on the blade tip surface of a large, 1.0 meter-chord, low-speed cascade representative of a high-pressure turbine. Data was obtained at a Reynolds number of 4.0 × 105 based on exit velocity and blade axial chord. Tip clearance levels ranged from 0.56% to 1.68% design span or equally from 1% to 3% of blade chord. An infrared camera, looking through the hollow blade, made detailed temperature measurements on a constant heat flux tip surface. The relative motion between the endwall and the blade tip was simulated by a moving belt. The moving belt endwall significantly to shifts the region of high Nusselt number distribution and reduces the overall averaged Nusselt number on the tip surface by up to 13.3%. The addition of a suction side squealer tip significantly reduced local tip heat transfer and resulted in a 32% reduction in averaged Nusselt number. Analysis of pressure measurements on the blade airfoil surface and tip surface along with PIV velocity flow fields in the gap give an understanding of the heat transfer mechanism.

Commentary by Dr. Valentin Fuster
2006;():483-490. doi:10.1115/GT2006-90480.

Compact heat exchanger designs are commonly used in many gas turbine engine applications. Though effective in their heat transfer function, they are often heavy, costly, and poor aerodynamic performers causing a reduction in engine efficiency. In addition, they are complex to manufacture and often prone to leakage. Finned surface heat exchangers are an attractive alternative to traditional compact designs. They can perform efficiently both aerodynamically and thermally. Such units could be mounted in the bypass fan stream of a gas turbine engine where large amounts of heat must be rejected from vital engine fluids such as oil and fuel. This research project investigated the efficiency of various fin designs applied to an oil cooler. Highly conductive materials, such as carbon composites were explored, and then compared to aerospace-quality aluminum alloys. Thermal, aerodynamic, economic, and weight performance comparisons between the carbon and aluminum fin structures were quantified. A three-dimensional numerical estimation of the final design concept was conducted using ANSYS. This research project specifically investigated the design of a finned surface air-oil heat exchanger. Design parameters included a total heat rejection of 2000 Btu/min and an oil temperature change of 100 degrees Fahrenheit with an inlet oil temperature of 300 degrees. The first design phase was conducted using an aerospace quality aluminum alloy. Internal and external flow convection theory was studied closely as well as basic heat exchanger and fin design concepts. A heat exchanger program was developed in Excel, automating the heat transfer based on basic geometric inputs. The program allowed easy iterations of fin/oil passage designs to meet the performance requirements and optimize the heat exchanger’s weight. The final iteration was then numerically modeled in ANSYS. The predicted heat transfer rate was then compared to the numerical estimation in ANSYS. The Excel program was validated by producing results within 2% of the ANSYS predicted solutions. Upon completion of the aluminum design. highly conductive materials, such as carbon composites were explored and implemented. The final designs of this project (both Aluminum and Carbon-Carbon) identified a new method of heat rejection at a significantly lower weight impact to the engine. The aluminum design had a total core weight of 25.4 lb while the carbon-carbon final design had a total core weight of 12.8 lb. In addition, both units have the potential to be incorporated within an existing engine case exposed to the bypass air stream, which may result in an additional weight savings.

Commentary by Dr. Valentin Fuster
2006;():491-501. doi:10.1115/GT2006-90516.

Gas turbine combustion chambers and turbine blades require better cooling techniques to cope with the increase in operating temperatures with each new engine model. Current gas turbine inlet temperatures are approaching 2000 K. Such extreme temperatures, combined with a highly dynamic environment, result in major stress on components, especially combustion chamber and blades of the first turbine stages. A technique that has been extensively investigated is transpiration cooling, for both combustion chambers and turbine blades. Transpiration-cooled components have proved an effective way to achieve high temperatures and erosion resistance for gas turbines operating in aggressive environments, though there is a shortage of durable and proven technical solutions. Effusion cooling (full-coverage discrete hole film cooling), on the other hand, is a relatively simpler and more reliable technique offering a continuous coverage of cooling air over the component’s hot surfaces. This paper presents an innovative technology for the efficient effusion cooling of the combustor wall and turbine blades. The dedicated electroforming process used to manufacture effusive film cooling systems, called Poroform®, is illustrated. A numerical model is also presented, developed specifically for designing the distributions of the diameter and density of the holes on the cooled surface with a view to reducing the metal’s working temperature and achieving isothermal conditions for large blade areas. Numerical simulations were used to design the effusive cooling system for a first-stage gas turbine blade. The diameter, density and spacing of the holes, and the adiabatic film efficiency are discussed extensively to highlight the cooling capacity of the effusive system.

Commentary by Dr. Valentin Fuster
2006;():503-512. doi:10.1115/GT2006-90532.

Effusion cooling of combustor liners for gas turbine engines is quite challenging and necessary to prevent thermal distress of the combustor liner walls. The flow and thermal patterns in the cooling layer are affected by the closely spaced film-cooling holes. It is important to fully document how the film layer behaves with a full-coverage cooling scheme to gain an understanding into surface cooling phenomena. This paper discusses experimental results from a combustor simulator tested in a low-speed wind tunnel. Engine representative, non-dimensional coolant flows were tested for a full-coverage effusion plate. Laser Doppler velocimetry was used to measure the flow characteristics of the cooling layer. These experiments indicate that the full-coverage film cooling flow has unique and scaleable velocity profiles that result from the closely spaced effusion holes. A parametric study of the cooling flow behavior illustrates the complex nature of the film flow and how it affects cooling performance.

Commentary by Dr. Valentin Fuster
2006;():513-521. doi:10.1115/GT2006-90534.

In a modern gas turbine engine the outer casing (shroud) of the shroudless high-pressure turbine is exposed to a combination of high flow temperatures and heat transfer coefficients. The casing is consequently subjected to high levels of convective heat transfer, a situation that is complicated by flow unsteadiness caused by periodic blade-passing events. In order to arrive at an over-tip casing design that has an acceptable service life it is essential for manfacturers to have appropriate predictive methods and cooling system configurations. It is known that both the flow temperature and boundary layer conductance on the casing wall vary during the blade-passing cycle. The current article reports the measurement of spatially and temporally resolved heat transfer coefficient (h) on the over-tip casing wall of a fully-scaled transonic turbine stage experiment. The results indicate that h is a maximum when a blade-tip is immediately above the point in question, while lower values of h are observed when the point is exposed to the rotor passage flow. Time-resolved measurements of static pressure are used to reveal the unsteady aerodynamic situation adjacent to the over-tip casing wall. The data obtained from this fully-scaled transonic turbine stage experiment are compared to previously published heat transfer data obtained in low-Mach number cascade style tests of similar high pressure blade geometries.

Commentary by Dr. Valentin Fuster
2006;():523-533. doi:10.1115/GT2006-90575.

This work supports new gas turbine designs for improved performance by evaluating endwall heat transfer rates in a cascade that is representative of a first stage stator passage and incorporates endwall assembly features and leakage. Assembly features, such as gaps in the endwall and misalignment of those gaps, disrupt the endwall boundary layer, typically leading to enhanced heat transfer rates. Leakage flows through such gaps within the passage can also affect endwall boundary layers and may induce additional secondary flows and vortex structures in the passage near the endwall. The present paper documents leakage flow and misalignment effects on the endwall heat transfer coefficients within a passage which has one axially-contoured and one straight endwall. In particular, features associated with the combustor-to-turbine transition piece and the assembly joint on the vane platform are addressed.

Commentary by Dr. Valentin Fuster
2006;():535-546. doi:10.1115/GT2006-90576.

The first stage vane section of a modern gas turbine engine is assembled with a gap between the combustor and the vane platform and a second gap on the platform. To prevent ingression, leakage flow is provided through each gap. In this paper, the effectiveness of the leakage flow as an endwall film coolant is measured. The cascade geometry includes axial contouring of the cooled endwall and several step configurations for each gap. The steps reflect assembly or differential thermal growth misalignment. Various blowing rates are applied through each gap to allow assessment of the changes in effectiveness with changes in leakage rate. Thus, the results presented herein show how the gaps, steps, and leakage rates alter the cooling effectiveness of the leakage flow. Shown are some cases where steps improve the film cooling effectiveness. In other cases, enhanced mixing due to gaps, steps, or increased leakage reduces effectiveness.

Commentary by Dr. Valentin Fuster
2006;():547-556. doi:10.1115/GT2006-90577.

The effects of obstructions on film cooling performance on a scaled-up 1st stage turbine vane will be discussed. Experimental results show that obstructions located upstream or inside of a film cooling hole will degrade adiabatic effectiveness up to 80% of the levels found with no obstructions. Downstream obstructions had little effect on performance. The location where the upstream obstructions ceased to degrade adiabatic effectiveness was determined and temperature profiles were constructed to determine how the upstream obstructions were affecting the mainstream and coolant flow.

Commentary by Dr. Valentin Fuster
2006;():557-567. doi:10.1115/GT2006-90581.

Oil film interferometry (OFI) is a measurement technique used to evaluate local wall shear stress and limiting streamline direction. It has been applied to aircraft wings in wind tunnel testing and in turbine cascade testing. The accuracy of the technique is largely dependent on the analysis of interferograms. A new technique for interferogram analysis is presented that is simple and particularly well suited to analyzing interferograms characterized by few fringes. For example, limited fringe definition can occur from dust contamination on the pressure surface of turbine airfoils during cascade wind tunnel testing. The new technique is described and illustrated with examples.

Commentary by Dr. Valentin Fuster
2006;():569-579. doi:10.1115/GT2006-90612.

This paper presents the experimental investigations on the discharge coefficients of impingement/effusion double flat wall with the impingement/effusion hole-area ratios of 1, 0.64 and 0.28. The impingement and effusion walls researched have equal numbers of holes per unit surface area in diamond arrays. The impingement holes are normal to the wall surface, the effusion holes are 30 degree to the wall surface in the sreamwise direction of the main flow, and both the impingement and effusion holes are arranged in the staggered mode. The CFD code was also applied to investigate the flow field within the impingement/effusion wall in detail. The experimental results indicated the relationship of the discharge coefficients of the single impingement wall, effusion wall and double wall with the overall pressure parameter of double wall in different impingement/effusion hole-area ratios, and were explained in the CFD results.

Commentary by Dr. Valentin Fuster
2006;():581-597. doi:10.1115/GT2006-90628.

Limited available data suggest a substantial impact of Mach number on the heat transfer from an array of jets impinging on a surface at fixed Reynolds number. Many jet array heat transfer correlations currently in use are based upon tests in which the jet Reynolds number was varied by varying the jet Mach number. Hence, this data may be inaccurate for high Mach numbers. Results from the present study are new and innovative because they separate the effects of jet Reynolds number and jet Mach number for the purposes of validating and improving correlations which are currently in use. The present study provides new data on the separate effects of Reynolds number and Mach number for an array of impinging jets in the form of discharge coefficients, local and spatially-averaged Nusselt numbers, and local and spatially-averaged recovery factors. The data are unique because data are given for impingement jet Mach numbers as high as 0.60 and impingement jet Reynolds numbers as high as 60,000, and because the effects of Reynolds number and Mach number are separated by providing data at constant Reynolds number as the Mach number is varied, and data at constant Mach number as the Reynolds number is varied. As such, the present data are given for experimental conditions not previously examined, which are outside the range of applicability of current correlations.

Commentary by Dr. Valentin Fuster
2006;():599-609. doi:10.1115/GT2006-90684.

The present paper reports on the aero-thermal performance of a nozzle vane cascade, with film cooled endwalls. The coolant is injected through four rows of cylindrical holes with conical expanded exits. Two endwall geometries with different area ratios have been compared. Tests have been carried out at low speed (M = 0.2), with coolant to mainstream mass flow ratio varied in the range 0.5–2.5%. Secondary flow assessment has been performed through 3D aerodynamic measurements, by means of a miniaturized 5-hole probe. Adiabatic effectiveness distributions have been determined by using the wide banded thermochromic liquid crystals (TLC) technique. For both configurations and for all the blowing conditions, the coolant share among the four rows has been determined. The aerothermal performance of the cooled vane have been analyzed on the basis of secondary flow effects and laterally averaged effectiveness distributions; this analysis was carried out for different coolant mass flow ratios. It was found that the smaller area ratio provides better results in terms of 3D losses and secondary flow effects; the reason is that the higher momentum of the coolant flow is going to better reduce the secondary flow development. The increase of the fan-shaped hole area ratio gives rise to a better coolant lateral spreading, but appreciable improvements of the adiabatic effectiveness were detected only in some regions and for large injection rates.

Commentary by Dr. Valentin Fuster
2006;():611-621. doi:10.1115/GT2006-90742.

Air film cooling has been successfully used to cool gas turbine hot sections for the last half century. A promising technology is proposed to enhance air film cooling with water mist injection. Numerical simulations have shown that injecting a small amount of water droplets into the cooling air improves film-cooling performance significantly. However, previous studies were conducted at conditions of low Reynolds number, temperature, and pressure to allow comparisons with experimental data. As a continuous effort to develop a realistic mist film cooling scheme, this paper focuses on simulating mist film cooling under typical gas turbine operating conditions of high temperature and pressure. The mainstream flow is at 15 atm with a temperature of 1561K. Both 2-D and 3-D cases are considered with different hole geometries on a flat surface, including a 2-D slot, a simple round hole, a compound-angle hole, and fan-shaped holes. The results show that 10%–20% mist (based on the coolant mass flow rate) achieves 5%–10% cooling enhancement and provides an additional 30–68K adiabatic wall temperature reduction. Uniform droplets of 5 to 20 μm are used. The droplet trajectories indicate the droplets tend to move away from the wall, which results in a lower cooling enhancement than under low pressure and temperature conditions. The commercial software Fluent (v. 6.2.16) is adopted in this study, and the standard k-ε model with enhanced wall treatment is adopted as the turbulence model.

Commentary by Dr. Valentin Fuster
2006;():623-632. doi:10.1115/GT2006-90784.

The requirements of always higher thermal performances of turbine vanes and blades push all the aeroengine industries to invest in new cooling technologies. AVIO, the italian aerospace propulsion industry, developed during the last three years an innovative cooling systems for stationary components, based on the doublewall technology, that allow to enhance the internal heat exchange optimising the cooling air distribution. The thermal behaviour of the system has been studied numerically using the standard integrated design methodologies. Some components have been then manufactured in real geometries and finally tested at high temperature in the AVIO’s burner rig. Aim of this work is to present the cooling design methodology used, the manufacturing phases, and to compare the measured thermal efficiency with the theoretical ones. The experimental tests of the blade at high temperature burner rig allowed to determine the effectiveness of the cooling system and validate the integrated aero-thermal numerical procedures. Experimental values of effectiveness are reported in relation with the numerical expected data. Therefore a comparison between this new technology and standard ones based on impingements and turbulated channels is taken into account. Overall numerical results have been found in good agreement with the experimental data and indicate that doublewall cooling system can provide heat transfer enhancements and higher thermal performance if compared with conventional cooling methods with potential benefits on cooling reduction and streamwise uniformity in heat transfer coefficients. Further future improvements and developments are here preliminary proposed.

Commentary by Dr. Valentin Fuster
2006;():633-645. doi:10.1115/GT2006-90802.

Internal cooling schemes for blades in a gas turbine engine often are subject to compromises between increased pressure losses in return for greater levels of heat transfer required to maintain durability levels in the engine’s harsh environment. Rib configurations have been the subject of much study in past years, however these configurations are normally presumed to be used in “full-coverage” mode, meaning that the ribs are placed in the channel in a continuous and uniform manner. This study investigates the interaction between the bend effects downstream of a 180° bend, which cause higher local heat transfer, and the effect of ribs. Some of the ribs directly downstream of the 180° bend in the 2nd leg of a two pass high aspect ratio (4:1) channel were removed and the effect on heat transfer was assessed. Experimental results showed that the heat transfer level recovered quickly once ribs were encountered. As expected, some decrease in heat transfer was observed in the region where ribs were removed; however total pressure losses in the channel were also much lower. Results include detailed two-dimensional heat transfer distributions determined by the transient liquid crystal method as well as an analysis of the balance between pressure recovery and local heat transfer levels. Generally, for the accuracy of the transient liquid crystal technique in complex three-dimensional flows, strongly varying fluid temperatures present in and downstream of the bend region must be taken into account. For this study, time and position dependent fluid temperature distributions were measured to account for these effects, making it possible to obtain high quality heat transfer results in those regions.

Commentary by Dr. Valentin Fuster
2006;():647-655. doi:10.1115/GT2006-90846.

This paper reports an experimental study of the thermal development in an idealized model of a blade cooling passage of smooth inner surfaces, comprising a square-ended U-bend with a cross-section that changes from a square upstream to a 2:1 rectangle downstream of the turn. The two flat walls are heated electrically, while the outer wall and the splitter plate are thermally insulated. The steady state liquid crystal technique is used to map the local Nusselt number variation. Measurements are obtained using a stationary air flow facility and also a rotating water flow facility. This enables us to investigate the effects on the thermal development of Reynolds variation from 30,000 to 100,000, Prandtl numbers of 0.7 and 5.8, and rotation numbers, from 0 to 0.4. The effects of minor modifications in the cross-sectional area at the bend exit, on the thermal development, under both stationary and rotating conditions, are also explored.

Commentary by Dr. Valentin Fuster
2006;():657-667. doi:10.1115/GT2006-90851.

A computational investigation is carried out to study the flow and heat transfer from a row of circular jets impinging on a concave surface. The computational domain simulates the impingement cooling zone of a gas turbine nozzle guide vane. The parameters which are varied in the study include jet Reynolds number (Red = 5000 to 54000), inter-jet distance to jet diameter ratio (c/d = 3.33 and 4.67) and target plate distance to jet diameter ratio (h/d = 1, 3 and 4). The flow field, predicated with K- ω turbulence model (using Fluent 6.2), is characterized with the presence of a pair of counter rotating vortices, an upwash fountain flow and entrainment. The local pressure coefficient and Nusselt number variations along the concave plate are presented and these values are found to under predict the some available experimental data by about 12%.

Commentary by Dr. Valentin Fuster
2006;():669-676. doi:10.1115/GT2006-90852.

In this study, a CFD-based optimization process is used to change the contour of the airfoil near a suction side cooling hole in order to improve its film effectiveness characteristics. An overview of the optimization process, which includes automated geometry, grid generation and CFD analyses is provided. From the results for the optimized geometry it is clear that the detachment of the cooling jet is much reduced and the cooling jet spread in the spanwise direction is increased substantially. The new external contour was then tested in a low-speed wind tunnel to provide a direct measure of the predictive capability. Comparisons to verification test data indicate that good agreement was achieved for both pressure and film cooling effectiveness behavior. This study proves that despite its limitations, current RANS methodology can be used a viable design tool and lead to innovative concepts for improving film cooling effectiveness.

Topics: Cooling , Airfoils
Commentary by Dr. Valentin Fuster
2006;():677-687. doi:10.1115/GT2006-90854.

Film-cooling in gas turbines leads to aerodynamic mixing losses and reduced temperatures of the gas flow. Improvements of the gas turbine thermal efficiency can be achieved by reducing the cooling fluid amount and by establishing a more equal distribution of the cooling fluid along the surface. It is well known that vortex systems in the cooling jets are the origin of reduced film-cooling effectiveness. For the streamwise ejection case, kidney-vortices result in a lift-off of the cooling jets; for the lateral ejection case, usually only one dominating vortex remains, leading to hot gas flow underneath the jet from one side. Based on the results of numerical analyses, a new cooling technology has been introduced by the authors, which reaches high film-cooling effectiveness as a result of a well-designed cooling hole arrangement for interaction of two neighbouring cooling jets (Double-jet Film-cooling DJFC). The results show that configurations exist, where an improved film-cooling effectiveness can be reached because an anti-kidney vortex pair is established in the double-jet. The paper aims on following major contributions: • to introduce the Double-jet Film-cooling (DJFC) as an alternative film-cooling technology to conventional film-cooling design. • to explain the major phenomena, which lead to the improvement of the film-cooling effectiveness by application of the DJFC. • to prove basic applicability of the DJFC to a realistic blade cooling configuration and present first test results under machine operating conditions.

Topics: Cooling
Commentary by Dr. Valentin Fuster
2006;():689-699. doi:10.1115/GT2006-90856.

A full-coverage cooled multi-layer plate configuration is investigated numerically by application of a 3-D conjugate fluid flow and heat transfer solver, CHTflow. The geometrical setup and the fluid flow conditions derive from modern gas turbine combustion chambers and bladings. The numerical grid contains the coolant supply (plenum), the solid body and the main flow area on the plate. The full-coverage cooling is realized by finest drilled holes with a diameter of 0.2 mm that are shaped in the region of the thermal barrier coating. Two different effects will be considered in this paper. First, the effects of deviations in the cooling hole geometry due to the manufacturing process will be discussed, then the effects of an obstruction in one of the cooling holes will be looked at. The numerical investigation focuses on the cooling efficiency on the plate surface and on the cooling hole surface as well as the aerodynamics in the main flow. The scientific work is embedded in the Collaborative Research Center 561 at RWTH Aachen University.

Commentary by Dr. Valentin Fuster
2006;():701-710. doi:10.1115/GT2006-90860.

The capabilities of four two-equation turbulence models in predicting film cooling effectiveness were investigated and their limitations as well as relative performance are presented. The four turbulence models are the standard, RNG, and realizable k-ε models as well as the standard k-ω model all found in the FLUENT CFD code. In all four models, the enhanced wall treatment has been used to resolve the flow near solid boundaries. A systematic approach has been followed in the computational setup to insure grid-independence and accurate solution that reflects the true capabilities of the turbulence models. Exact geometrical and flow-field replicas of an experimental study on discrete-jet film cooling were generated and used in FLUENT. A pitch-to-diameter ratio of 3.04, injection length-to-diameter ratio of 4.6 and density ratios of 0.92 and 0.97 were some of the parameters used in the film cooling analysis. Furthermore, the study covered two levels of blowing ratio (M = 0.5 and 1.5) at an environment of low free-stream turbulence intensity (Tu = 0.1%). The standard k-ε model had the most consistent performance among all considered turbulence models and the best centerline film cooling effectiveness predictions with the results deviating from experimental data by only ±10% and about 20–60% for the low (M = 0.5) and high (M = 1.5) blowing ratio cases, respectively. However, centerline side-view and surface top-view contours of non-dimensional temperature for the standard k-ε cases revealed that the good results for film cooling effectiveness η compared to the experimental data were due to a combination of an over-prediction of jet penetration in the normal direction with an under-prediction of jet spread in the lateral direction. The standard k-ω model completely failed to produce any results that were meaningful with under-predictions of η that ranged between 80 and 85% for the low blowing ratio case and over-predictions of about 200% for the high blowing ratio case. Even though the RNG and realizable models showed to have better predicted the jet spread in the lateral direction compared to the standard k-ε model, there were some aspects of the flow, such as levels of turbulence generated by cross-flow and jet interaction, that were not realistic resulting in errors in the η prediction that ranged from −10% to +80% for the M = 0.5 case and from −80% to +70% for the M = 1.5 case. As a result of this study at this point it was concluded that the standard k-ε model have the most promising potential among the two-equation models considered. It was chosen as the best candidate for further improvement for the simulation of film cooling flows.

Commentary by Dr. Valentin Fuster
2006;():711-720. doi:10.1115/GT2006-90866.

Experimental investigation was carried out to study the heat transfer characteristics in a rectangular duct cooled by an array of impinging jets. Air ejected from an array of orifices is aimed at the heated target surface and exits from the radial outlets. The effect of feed channel widths (5 ≤ H / d ≤ 9), jet to target plate distance (4 ≤ S/d ≤ 8), outflow orientation and jet Reynolds numbers (9300 ≤ Rej ≤ 18800) with a single array of equally spaced off-set orifice jets of diameter d = 0.5 cm on heat transfer was studied. Results indicated that the outflow orientation causing crossflow effect significantly affects the Nusselt number distributions on the target surface. Relatively higher Nusselt number values were obtained for the outflow orientation where the flow exits in both the directions. The feed channel width H/d = 7 gave relatively higher values of heat transfer compared to the other two feed channel widths. The jet-to-plate distance S/d = 4 resulted in higher heat transfer compared to the other jet-to-plate distances.

Commentary by Dr. Valentin Fuster
2006;():721-737. doi:10.1115/GT2006-90927.

This paper presents both measurements and predictions of the hot-gas-side heat transfer to a modern, one and 1/2 stage high-pressure, transonic turbine. Comparisons of the predicted and measured heat transfer are presented for each airfoil at three locations, as well as on the various endwalls and rotor tip. The measurements were performed using the Ohio State University Gas Turbine Laboratory Test Facility (TTF). The research program utilized an uncooled turbine stage at a range of operating conditions representative of the engine: in terms of corrected speed, flow function, stage pressure ratio, and gas-to-metal temperature ratio. All three airfoils were heavily instrumented for both pressure and heat transfer measurements at multiple locations. A 3-D, compressible, Reynolds-averaged Navier-Stokes CFD solver with k-ω turbulence modeling was used for the CFD predictions. The entire, 1-1/2 stage turbine was solved using a single computation, at two different Reynolds numbers. The CFD solutions were steady, with tangentially mass-averaged inlet/exit boundary condition profiles exchanged between adjacent airfoil-rows. Overall, the CFD heat transfer predictions compared very favorably with both the global operation of the turbine and with the local measurements of heat transfer. A discussion of the features of the turbine heat transfer distributions, and their association with the corresponding flow-physics, has been included.

Commentary by Dr. Valentin Fuster
2006;():739-750. doi:10.1115/GT2006-90949.

A new, computationally-efficient method is presented for processing transient thin-film heat transfer gauge signals. These gauges are widely used in gas turbine heat transfer research, where, historically, the desired experimental heat transfer flux signals, q, are derived from transient measured surface-temperature signals, T, using numerical approximations to the solutions of the linear differential equations relating the two. The new method uses known pairs of exact solutions, such as the T response due to a step in q, to derive a sampled approximation of the impulse response of the gauge system. This impulse response is then used as a finite impulse response digital filter to process the sampled T signal to derive the required sampled q signal. This is computationally efficient because the impulse response need only be derived once for each gauge for a given sample rate, but can be re-used repeatedly, using optimised Matlab filter routines and is highly accurate. The impulse response method can be used for most types of heat flux gauge. In fact, the method is universal for any linear measurement systems which can be described by linear differential equations where theoretical solution pairs exist between input and output. Examples using the new method to process turbomachinery heat flux signals are given.

Commentary by Dr. Valentin Fuster
2006;():751-762. doi:10.1115/GT2006-90966.

This paper describes the experimental approach utilized to perform experiments using a fully cooled rotating turbine stage to obtain film effectiveness measurements. Significant changes to the previous experimental apparatus were implemented to meet the experimental objectives. The modifications include the development of a synchronized blowdown facility to provide cooling gas to the turbine stage, installation of a heat exchanger capable of generating a uniform or patterned inlet temperature profile, novel utilization of temperature and pressure instrumentation, and development of robust double-sided heat flux gauges. With these modifications, time-averaged and time-accurate measurements of temperature, pressure, surface heat flux, and film effectiveness can be made over a wide range of operational parameters duplicating the non-dimensional parameters necessary to simulate engine conditions. Data from low Reynolds number experiments are presented to demonstrate that all appropriate scaling parameters can be satisfied and that the new components have operated correctly. Along with airfoil surface heat transfer and pressure data, temperature and pressure data from inside the coolant plenums of the vane and rotating blade airfoils are presented. Pressure measurements obtained inside the vane and blade plenum chambers illustrate passing of the wakes and shocks as a result of vane/blade interaction. Part II of this paper presents data from the low Reynolds number cooling experiments and compares these measurements to CFD predictions generated using the Numeca FINE/Turbo package at multiple spans on the vanes and blades.

Commentary by Dr. Valentin Fuster
2006;():763-776. doi:10.1115/GT2006-90968.

This paper presents measurements and the companion CFD predictions for a fully cooled, high-work single stage HP turbine operating in a short-duration blowdown rig. Part I of this paper presented the experimental approach, and Part II focuses on the results of the measurements and demonstrates how these results compare to predictions made using the Numeca FINE/Turbo CFD package. The measurements are presented in both time-averaged and time-accurate formats. The results include the heat transfer at multiple spans on the vane, blade, and rotor shroud as well as flow path measurements of total temperature and total pressure. Surface pressure measurements are available on the vane at midspan, and on the blade at 50% and 90% spans as well as the rotor shroud. In addition, temperature and pressure measurements obtained inside the coolant cavities of both the vanes and blades are presented. Time-averaged values for the surface pressure on the vane and blade are compared to steady CFD predictions. Additional comparisons will be made between the heat transfer on cooled blades and uncooled blades with identical surface geometry. This, along with measurements of adiabatic wall temperature, will provide a basis for analyzing the effectiveness of the film-cooling scheme at a number of locations.

Commentary by Dr. Valentin Fuster
2006;():777-785. doi:10.1115/GT2006-90973.

The present research investigates the heat transfer characteristics in an equilateral triangular channel to simulate the leading edge cooling passage of a gas turbine blade. The experiments are conducted for the stationary and rotating ribbed channel with three different attack angles (45°, 90° and 135°). Square ribs are installed in a staggered manner on the pressure and suction side surfaces of the channel. The rib height to channel hydraulic diameter ratio (e/Dh ) is 0.079 and the rib-to-rib pitch (p) is 8 times of the rib height. To measure regional-averaged heat transfer coefficients in the channel, two rows of copper blocks with heaters are installed on each surface. The rotation number ranges from 0.0 to 0.1 for the fixed Reynolds number of 10,000. Inlet coolant-to-surface density ratio is about 0.2. For the channel with 90° ribs, the heat transfer rates of all regions have similar values for stationary case. However, for the rotating channel, heat transfer coefficients on the pressure side surface are significantly increased while the suction side surface has quite low heat transfer coefficients due to a single rotating secondary flow induced by Coriolis force. For the channel with angled rib arrangements, a pair of counter-rotating vortices is induced by the angled rib arrangements. High heat transfer coefficients are obtained on the regions near the inner wall for 45° angled ribbed channel and near the leading edge for the 135° angled ribbed channel. The heat transfer coefficients in rotating channel with angled ribs are almost the same as those of stationary case for the tested conditions because the secondary flow dominates the heat transfer. The channel with angled ribs consistently yields better thermal performance than the transverse ribbed channel for the test conditions of the present study.

Commentary by Dr. Valentin Fuster
2006;():787-803. doi:10.1115/GT2006-91013.

Aircraft propulsion engines, land-based power generation, and industrial machines have, as a primary component, the turbine as means to produce thrust or generate power. In the turbine section of the engine, airfoil components are subjected to extremely complex and damaging environments. The combination of high gas temperatures and pressures, strong gradients, abrupt geometry changes, viscous forces, rotational forces, and unsteady turbine vane/blade interactions, all combine to offer a formidable challenge in terms of turbine durability. Nevertheless, the ultimate goal is to maintain or even improve the highest level of turbine performance and simultaneously reduce the amount of cooling flow needed to achieve this end. As such, coolant flow is a penalty to the cycle and thermal efficiency. Cooling strategies are developed and presented to determine ways for coolant flow management. The main variables include film cooling configurations, and convective efficiency schemes to balance turbine airfoil thermal loads for target overall cooling effectiveness. The desired targets are determined by the turbine airfoil durability requirements of oxidation and fatigue on a local scale and for creep on the bulk scale. Emphasis is provided to the general modes of cooling including film cooling, impingement cooling, and convective cooling for different parts of the airfoil such as leading edge, mid-body, trailing edge, tip and endwalls. Convective cooling is presented in terms of fundamental cooling enhancements, such as turbulating trip strips and pedestals. Recent literature dealing with these topics is listed.

Commentary by Dr. Valentin Fuster
2006;():805-817. doi:10.1115/GT2006-91017.

Modern gas turbine engines provide large amounts of thrust and experience severe cyclic thermal gradients with sustained mechanical loads. Fatigue damage induced by cyclic thermal transients, along with creep and oxidation damage caused by quasi steady state, long hold-time periods, have emerged as critical life-limiting factors for coated turbine airfoils. It is, therefore, necessary to develop key thermal parameters for life prediction system for hot section components that accounts for operating thermal transients of anisotropic material properties for coated airfoils. Thermal cyclic deformations are determined as a function of the cyclic temperatures and sustained mechanical loads. This leads to the evaluation of local cyclic fatigue damage function, which can be compared against the strain energy density dissipated during monotonic deformation testing for damage assessment. Long hold-time periods at elevated temperatures also lead to cumulative effects of creep and oxidation. Key thermal-mechanical parameters, obtained from the formulation presented, arise as governing parameters, leading to analytical maps necessary to track the damage observed in cyclic thermomechanical test specimens. The contributing damage modes of thermal fatigue, creep and oxidation are summed by the traditional Miner’s rule of accumulated damage. This provides a way to determine the number of cycles remaining in the part.

Commentary by Dr. Valentin Fuster
2006;():819-828. doi:10.1115/GT2006-91019.

Described in this paper is an experimental study of heat transfer over a trailing edge configuration preceded with an internal cooling channel of pedestal array. The pedestal array consists of both circular pedestals and oblong shaped blocks. Downstream to the pedestal array, the trailing edge features pressure side cutback partitioned by the oblong shaped blocks. The local heat transfer coefficient over the entire wetted surface in the internal cooling chamber has been determined by using a “hybrid” measurement technique based on transient liquid crystal imaging. The hybrid technique employs the transient conduction model in a semi-infinite solid for resolving the heat transfer coefficient on the endwall surface uncovered by the pedestals. The heat transfer coefficient over a pedestal can be resolved by the lumped capacitance method with an assumption of low Biot number. The overall heat transfer for both the pedestals and endwalls combined shows a significant enhancement compared to the case with thermally developed smooth channel. Near the downstream most section of the suction side, the land, due to pressure side cutback, is exposed to the stream mixed with hot gas and discharged coolant. Both the adiabatic effectiveness and heat transfer coefficient on the land section are characterized by using the transient liquid crystal technique.

Commentary by Dr. Valentin Fuster
2006;():829-838. doi:10.1115/GT2006-91039.

A method allowing the evaluation of the effects related to heat transfer to the turbine blades on its performance characteristics is presented. The effects investigated are the change of passage dimensions, resulting from heat transfer and the change in flow field, exhibited mainly as a different boundary layer development. Change of hot gas temperature combined with cooling air temperature and possibly flow rate, result in a change of the temperature of the blade material, leading to dimension changes, because of the thermal expansion (dilatation). The changes in dimensions have a direct effect on turbine performance. An immediate consequence is a modification of the mass flow characteristic, due to a change of the throat area. Heat transfer also influences the properties of the gas flowing through the passage and in particular the characteristics of the boundary layers developing on the nozzle vanes and hub, tip endwals. Change of the thickness of this layer results in a change of blockage through the passage, a fact that influences directly the turbine flow function. The influence of both effects on turbine performance is studied. The study is performance oriented, aiming to the derivation of simplified models, which can be introduced in engine cycle decks.

Commentary by Dr. Valentin Fuster
2006;():839-848. doi:10.1115/GT2006-91067.

Currently, turbines are being designed to operate with increasing inlet temperatures to improve the engine’s performance. To reduce NOx combustion, combustors are being designed to provide flat pattern factors. For these reasons, the endwall of the first stage vane is under severe heat transfer conditions. Film-cooling is one of the most effective cooling methods for the endwall of the vane. This paper presents results from a computational study of a film-cooled endwall. The endwall design considers both an upstream slot and a mid-passage slot, whereby the slot considered is both aligned and misaligned with respect to the endwall. Results indicate reasonable agreement between computational predictions and experimental measurements of adiabatic effectiveness levels along the vane endwall. The results of this study show the mid-passage slot has a large effect on the endwall film coverage. In addition, the relative height of the upstream slot to the downstream endwall is important to consider for improving the cooling benefit from the leakage flow between the combustor and turbine.

Topics: Cooling
Commentary by Dr. Valentin Fuster
2006;():849-863. doi:10.1115/GT2006-91073.

A feature-based jet model has been proposed for use in 3D CFD prediction of turbine blade film cooling. The goal of the model is to be able to perform computationally efficient flow prediction and optimization of film-cooled turbine blades. The model reproduces in the near hole region the macro flow features of a coolant jet within a Reynolds-Averaged Navier Stokes (RANS) framework. Numerical predictions of the 3D flow through a linear transonic film-cooled turbine cascade are carried out with the model, with a low computational overhead. Different cooling holes arrangement are computed and the prediction accuracy is evaluated versus experimental data. It shown that the present model provides a reasonably good prediction of the adiabatic film-cooling effectiveness and Nusselt number around the blade. A numerical analysis of the interaction of coolant jets issuing from different rows of holes on the blade pressure side is carried out. It is shown that the upward radial migration of the flow due to the passage secondary flow structure has an impact on the spreading of the coolant and the film cooling effectiveness on the blade surface. Based on this result, a new arrangement of the cooling holes for the present case is proposed that leads to a better spanwise covering of the coolant on the blade pressure side surface.

Commentary by Dr. Valentin Fuster
2006;():865-875. doi:10.1115/GT2006-91109.

The conjugate heat transfer methodology has been employed to predict the flow and thermal properties including the metal temperature of a NASA turbine vane at three operating conditions. The turbine vane was cooled internally by air flowing through 10 round pipes. The conjugate heat transfer methodology allows a simultaneous solution of aerodynamics and heat transfer in the external hot gas and the internal cooling passages and conduction within the solid metal, eliminating the need for multiple/decoupled solutions in a typical industry design process. The model of about three million computational meshes includes the gas path and the internal cooling channels, comprising hexa cells, and the solid metal comprising hexa and prism cells. The predicted aerodynamic loadings were found to be in close agreement with the data for all the cases. The predicted metal temperature, external and internal heat transfer distributions at the mid-span compared well with the measurement. The differences in the heat transfer rates and metal temperature under different running conditions were also captured well. The V2F turbulence model has been compared with a low-Reynolds-number k-ε model and a non-linear quadratic k-ε model. The V2F model is found to provide the closest agreement with the data, though it still has room for improvement in predicting the boundary layer transition and turbulent heat transfer, especially on the suction side. The overall results are quite encouraging and indicate that conjugate heat transfer simulation with proper turbulence closure has the potential to become a viable tool in turbine heat transfer analysis and cooling design.

Commentary by Dr. Valentin Fuster
2006;():877-885. doi:10.1115/GT2006-91122.

The present study investigates the effects of bleed flow on heat/mass transfer and pressure drop in a rotating channel with transverse rib turbulators. The hydraulic diameter (Dh ) of the square channel is 40.0 mm. The bleed holes are located between the rib turburators on the leading surface and the hole diameter (d) is 4.5 mm. The square rib turbulators are installed on both leading and trailing surface. The rib-to-rib pitch (p) is 10.0 times of the rib height (e) and the rib height-to-hydraulic diameter ratio (e/Dh ) is 0.055. The tests were conducted at various rotation numbers (0, 0.2, 0.4), while the Reynolds number and the rate of bleed flow to main flow were fixed at 10,000 and 10%, respectively. A naphthalene sublimation method was employed to determine the detailed local heat transfer coefficients using the heat/mass transfer analogy. The results suggest that for a rotating ribbed passage with bleed flow of BR = 0.1, the heat/mass transfer on the leading surface is dominantly affected by rib turbulators and the secondary flow induced by rotation rather than bleed flow. The heat/mass transfer on the trailing surface decreases due to the diminution of main flow. The results also show that the friction factor decreases with bleed flow.

Commentary by Dr. Valentin Fuster
2006;():887-898. doi:10.1115/GT2006-91124.

A numerical investigation is carried out for estimating the influence of rib turbulators on heat transfer and pressure drop of staggered non-uniform pin-fin arrays of different shapes, in a simulated cambered vane trailing region. Pin-fins of square, circular and the diamond shapes, each of two sizes (d) were chosen. The ratio of span-wise pitch to longitudinal pitch is 1.06 and that to the pin size are 4.25 and 3.03, for all pin shapes. A constant heat flux boundary condition is assumed over the coolant channel walls, rib surfaces and circumferential faces of the pin-fins. Reynolds number is varied (20,000<ReD <40,000) by changing the coolant outlet to inlet pressure ratio. Pin end-wall and pin surface averaged heat transfer coefficients and Nusselt numbers are estimated and detailed contours of heat transfer coefficient on both the pressure and suction surfaces are presented. Whilst there is an enhancement in heat transfer and pressure drop with ribs for all the pin shapes, diamond pins have shown the highest enhancement values for both ribbed and non-ribbed configuration.

Commentary by Dr. Valentin Fuster
2006;():899-908. doi:10.1115/GT2006-91129.

A novel heat flux sensor was tested which allows for time-resolved heat flux measurements in internal ribbed channels related to the study of passages in gas turbine blades. The working principle of the Atomic Layer Thermopile (ALTP) sensor is based on a thermoelectric field created by a temperature gradient over an YBCO crystal (the transverse Seebeck effect). The sensors very fast frequency response allows for highly time-resolved heat flux measurements up to the 1 MHz range. This paper explains the design and working principle of the sensor, as well as the benchmarking of the sensor for several flow conditions. For internal cooling passages, this novel sensor allows for highly accurate, time-resolved measurements of heat transfer coefficients, leading to a greater understanding of the influence of fluctuations in temperature fields.

Commentary by Dr. Valentin Fuster
2006;():909-920. doi:10.1115/GT2006-91131.

Measurements using a novel heat flux sensor were performed in an internal ribbed channel representing the internal cooling passages of a gas turbine blade. These measurements allowed for the characterization of heat transfer turbulence levels and unsteadiness not previously available for internal cooling channels. In the study of heat transfer, often the fluctuations can be equally as important as the mean values for understanding the heat loads in a system. In this study comparisons are made between the time-averaged values obtained using this sensor and detailed surface measurements using the transient thermal liquid crystal technique. The time-averaged heat flux sensor and transient TLC results showed very good agreement, validating both methods. Time-resolved measurements were also corroborated with hot film measurements at the wall at the location of the sensor to better clarify the influence of unsteadiness in the velocity field at the wall on fluctuations in the heat flux. These measurements resulted in turbulence intensities of the velocity and heat flux of about 20%. The velocity and heat flux integral length scales were about 60% and 35% of the channel width respectively, resulting in a turbulent Prandtl number of about 1.7 at the wall.

Commentary by Dr. Valentin Fuster
2006;():921-929. doi:10.1115/GT2006-91161.

Flow aligned blockers are proposed to minimize the entrainment of hot gases underneath film-cooling jets by the counter-rotating vortices within the jets. Computations, based on the ensemble-averaged Navier-Stokes equations closed by the realizable k-ε turbulence model, were used to assess the usefulness of rectangular prisms as blockers in increasing film-cooling adiabatic effectiveness without unduly increasing surface heat transfer and pressure loss. The Taguchi’s design of experiment method was used to investigate the effects of the height of the blocker (0.2D, 0.4D, 0.8D), the thickness of the blocker (D/20, D/10, D/5), and the spacing between the pair of blockers (0.8D, 1.0D, 1.2D), where D is the diameter of the film-cooling hole. The effects of blowing ratio (0.37, 0.5, 0.65) were also studied. Results obtained show that blockers can greatly increase film-cooling effectiveness. By using rectangular prisms as blockers, the laterally averaged adiabatic effectiveness at 15D downstream of the film-cooling hole is as high as that at 1D downstream. The surface heat transfer was found to increase slightly near the leading edge of the prisms, but reduced elsewhere from reduced temperature gradients that resulted from reduced hot gas entrainment. However, pressure loss was found to increase somewhat because of the flat rectangular leading edge, which can be made more streamlined.

Commentary by Dr. Valentin Fuster
2006;():931-938. doi:10.1115/GT2006-91163.

A new design concept is presented to increase the adiabatic effectiveness of film cooling jets without unduly increasing surface heat transfer and pressure loss. Instead of shaping the film-cooling hole at its downstream end as is done for shaped holes, this study proposes a geometry modification upstream of the film-cooling hole to modify the approaching boundary-layer flow and its interaction with the film-cooling jet. Computations, based on the ensemble-averaged Navier-Stokes equations closed by the realizable k-ε turbulence model, were used to examine the usefulness of making the surface just upstream of the film-cooling hole into a ramp with backward-facing step. The effects of the following parameters were investigated: angle of the ramp (8.5°, 10°, 14°), distance between the backward-facing step of the ramp and the film-cooling hole (0.5D, D), and blowing ratio (0.36, 0.49, 0.56, 0.98). Results obtained show that an upstream ramp with a backward-facing step can greatly increase film-cooling adiabatic effectiveness. The laterally averaged adiabatic effectiveness with ramp can be two or more times higher than without the ramp. Also, the ramp increases the surface area that each film-cooling jet protects. However, using the ramp does increase drag. The increase in surface heat transfer was found to be minimal.

Topics: Cooling
Commentary by Dr. Valentin Fuster
2006;():939-946. doi:10.1115/GT2006-91168.

Numerical and experimental approaches were taken to improve the heat transfer performance of V-shaped staggered (VSG) ribs. At first, a numerical method using Large Eddy Simulations (LES) was employed to determine the effect of VSG rib-induced flow on the local heat transfer coefficient distributions. The results revealed that the secondary flows generated by the rib configuration carry cold coolant air from the passage core region to the rib-roughened wall, thus enhancing heat transfer. Conversely, behind each rib there is a recirculation zone that does not contribute to enhanced heat transfer. Based on said results, six types of the advanced V-shaped staggered (AVSG) rib configurations were proposed. The test apparatus employing a comparative method was used to measure the total heat transfer coefficient and pressure loss coefficient of the VSG and AVSG ribs. It was concluded that thermal performance of the most effective type was 25% higher than that of VSG.

Commentary by Dr. Valentin Fuster
2006;():947-956. doi:10.1115/GT2006-91194.

One option to improve the cycle efficiency of current state-of-the-art aero engines is to increase the turbine inlet temperature. Since this temperature is above the melting temperature for the alloys utilised in the turbine component already today, efficient cooling methods must be developed that consider both aerodynamic and aerothermal aspects of cooling. Here, the goal is to extract as little as possible secondary air from the main hot gas cycle for cooling and to use this coolant then aerodynamically and aerothermally as efficient as possible. The paper to be presented documents a CFD based design approach that lead to a new passive shroud cooling concept and the definition of its operational parameters. By using a simple one dimensional method [10] for predicting the aerodynamic losses resulting from such a cooling configuration in connection with 3d Navier-Stokes solvers (RANS) for predicting film cooling effectiveness contours on the rotor shroud surfaces, the new cooling configuration was developed. The concept was then tested and confirmed experimentally as documented in more detail in Part 2 of this paper. It is noted that only 70% of the coolant mass flow required for the current configuration was used for the new concept whereas the aerodynamic efficiency measured remained nearly constant. Improving upon existing passive shroud cooling systems where the coolant is injected directly into the labyrinth of the shroud, the new approach comprises cooling holes that inject the coolant upstream of the labyrinth and through the stator platform into the main passage flow. Here, it is important that the bulk of the coolant is placed below the dividing streamlines between main passage flow and labyrinth flow. Thereby, it can be achieved that the major part of the coolant indeed reaches the thermally loaded target surfaces on the shroud bottom at various axial gaps due to different operating points of the turbine. Besides the improved film-cooling effectiveness measured, the second important aspect of the new concept is the achievement of as small as possible additional aerodynamic losses due to coolant ejection into a high speed flow region. It will be shown that both goals can be achieved by the new concept. Furthermore, CFD results on film-cooling performance and aerodynamic losses will be shown and compared with experimental data.

Topics: Cooling , Design , Turbines
Commentary by Dr. Valentin Fuster
2006;():957-967. doi:10.1115/GT2006-91201.

Proper cooling of the airfoil trailing edge is imperative in gas turbine designs since this area is often one of the life limiting areas of an airfoil. A common method of providing thermal protection to an airfoil trailing edge is by injecting a film of cooling air through slots located on the airfoil pressure side near the trailing edge, thereby providing a cooling buffer between the hot mainstream gas and the airfoil surface. In the conventional designs, at the breakout plane, a series of slots open to expanding tapered grooves in between the tapered lands and run the cooling air through the grooves to protect the trailing edge surface. In this study, naphthalene sublimation technique was used to measure area averaged mass/heat transfer coefficients downstream of the breakout plane on the slot and on the land surfaces. Three slot geometries were tested: a) a baseline case simulating a typical conventional slot and land design, b) same geometry with a sudden outward step at the breakout plane around the opening, and c) the sudden step was moved one-third away from the breakout plane in the slot. Mass/heat transfer results were compared for these slots geometries for a range of blowing ratios (M = (ρu)s /(ρu)m ) from 0 to 2. For the numerical investigation, a pressure-correction based, multi-block, multi-grid, unstructured/adaptive commercial software was used in this investigation. Several turbulence models including the standard high Reynolds number k-ε turbulence model in conjunction with the generalized wall function were used for turbulence closure. The applied thermal boundary conditions to the CFD models matched the test boundary conditions. Effects of a sudden downward step (Coanda) in the slot on mass/heat transfer coefficients on the slot and on the land surfaces were compared both experimentally and numerically.

Commentary by Dr. Valentin Fuster
2006;():969-981. doi:10.1115/GT2006-91202.

Reliance on Thermal Barrier Coatings (TBC) to reduce the amount of air used for turbine vane cooling is beneficial both from the standpoint of reduced NOx production, and as a means of improving cycle efficiency through improved component efficiency. It is shown that reducing vane cooling from 10% to 5% of mainstream air can lead to NOx reductions of nearly 25% while maintaining the same rotor, inlet temperature. An analysis is given which shows that, when a TBC is relied upon in the vane thermal design process, significantly less coolant is required using internal cooling alone compared to film cooling. This is especially true for small turbines where internal cooling without film cooling permits the surface boundary layer to remain laminar over a significant fraction of the vane surface.

Commentary by Dr. Valentin Fuster
2006;():983-993. doi:10.1115/GT2006-91207.

The present experimental study investigates the aero performance differences between a conventional turbine blade trailing edge and a trailing edge with a sharp cut-back. Both geometries include trailing edge film cooling. A scaled model of a conventional turbine blade trailing edge and the trailing edge with a sharp cut-back including the scaled film cooling hole geometries were incorporated into flat plates. Experiments were conducted in a low speed wind tunnel to establish the performance change caused by introducing such a cut-back system on the trailing edge. Experiments were conducted at a Reynolds number of 1.9 × 106 at the trailing edge and for blowing ratios from 0.9 to 2.4. The experimental data presented include; static pressure variation on the plate surfaces; the change in discharge coefficient due to the cut-back; detailed total pressure measurements via a 2D traverse mechanism at a plane 40mm downstream of the flat plates. The total pressure measurements were used to establish the mixed out loss for both configurations.

Commentary by Dr. Valentin Fuster
2006;():995-1004. doi:10.1115/GT2006-91229.

Impingement systems are common place in many turbine cooling applications. Generally these systems consist of a target plate that is cooled by the impingement of multiple orthogonal jets. While it is possible to achieve high target surface heat transfer with this configuration, the associated pressure drop is generally high and the cooling efficiency low. Furthermore, especially in large impingement arrays, the build-up of cross flow from upstream jets can be significant and result in deflection of downstream impingement jets reducing the resultant heat transfer coefficient distribution. This paper presents a computational and experimental investigation into the use of shaped elliptical or elongated circular impingement holes designed to improve the penetration of the impinging jet across the coolant passage. This is of particular interest where there is significant cross flow. Literature review and computational investigations are used to determine the optimum aspect ratio of the impingement jet. The improved heat transfer performance of the modified design is then tested in an experimental rig with varying degrees of cross flow at engine representative conditions. In all cases a 16% increase in the Nusselt number on the impingement target surface in the downstream half of the cooling passage was achieved. Under the first 4 impingement holes Nusselt number enhancement of enhancement of 28–77% was achieved provided no additional cross flow was present in the passage. When appropriately aligned, a significant reduction in the stress concentration factor caused by the addition of a hole can be achieved using this design.

Commentary by Dr. Valentin Fuster
2006;():1005-1014. doi:10.1115/GT2006-91230.

Naphthalene sublimation experiments were conducted to study heat transfer for flow through blockages with holes in an internal cooling passage near the trailing edge of a gas turbine airfoil. The cooling passage was modeled as two rectangular channels whose heights decreased along the main flow direction. The air made a right-angled turn before passing through two blockages with staggered holes in each channel, and left the channel through an exit that was partially blocked by periodic lands with rounded leading edges. There were ten holes along each blockage and all of the holes had rounded edges. Local heat (mass) transfer was measured, and overall heat (mass) transfer results were obtained, on the exposed surfaces of one of the walls downstream of the two blockages, for Reynolds numbers (based on the hydraulic diameter of the channel at the upstream surface of the first blockage) between 5,000 and 36,000. The results showed that the blockages with the larger hole-to-channel cross-sectional area ratio in one of the two test sections enhanced the heat (mass) transfer downstream of the blockages more than the blockages with the smaller open area ratio in the second test section. For the geometric configurations and flow conditions studied, the average heat (mass) transfer was higher downstream of the second blockage than downstream of the first blockage. The configurations of the inlet channel and the exit slots considered in this study did not significantly affect the local heat (mass) transfer distributions or the average heat (mass) transfer downstream of the blockages.

Commentary by Dr. Valentin Fuster
2006;():1015-1026. doi:10.1115/GT2006-91231.

Detailed measurements of the heat transfer coefficient distributions on the internal surfaces of a novel gas turbine blade cooling configuration were carried out using a transient liquid crystal technique. The cooling geometry, in which a series of racetrack passages are connected to a central plenum, provides high heat transfer coefficients in regions of the blade in good thermal contact with the outer blade surface. The Reynolds number changes along its length because of the ejection of fluid through a series of 19 transfer holes in a staggered arrangement, which are used to connect ceramic cores during the casting process. Heat transfer coefficient distributions on this holes surface are particularly important in the prediction of blade life, as are heat transfer coefficients within the hole. Results at passage inlet Reynolds numbers of 21667, 45596 and 69959 are presented along with in-hole htc distributions at Rehole = 5930, 12479, 19147 and suction ratios of 0.98, 1.31, 2.08, 18.67. All values are engine representative. The results were compared to predictions made using the commercial CFD package Fluent. Characteristic regions of high heat transfer downstream of the transfer holes were observed with enhancement of up to 92% over the Dittus-Boelter level. Within the transfer holes, the average htc level was strongly affected by the crossflow at the hole entrance. Htc levels were low in these short (l/d = 1.5) holes fed from regions of developed boundary layer.

Commentary by Dr. Valentin Fuster
2006;():1027-1039. doi:10.1115/GT2006-91237.

In this paper detailed experimental measurements and computational predictions of heat transfer coefficient distributions in a large scale perspex model of a novel integrally cast blade cooling geometry are reported. In a gas turbine blade, the cooling passage investigated is integrally cast into the blade wall, providing good thermal contact with the outer surface of the turbine blade. Flow enters the racetrack passage through the root of the blade and exits to a central plenum through a series of nineteen transfer holes equally spaced in a staggered arrangement across the span of the blade. The Reynolds number changes continuously along the passage length because of the continuous ejection of fluid through a series of 19 transfer holes to the plenum. The smooth passage surface opposite is in closest proximity to the external surface, and this investigation has characterised the heat transfer coefficient on this surface at a range of engine representative inlet Reynolds numbers using a hybrid transient liquid crystal technique. The ability of three different rib configurations to enhance the heat transfer on this surface was also determined. Because the passage at engine scale is necessarily small, the rib height in all cases was 32.5% of the passage height. As the entire passage wetted surface is able to contribute to the blade cooling, and knowledge of the heat transfer coefficient distribution on the holed surfaces is crucial to prediction of blade life, a commercial CFD package, Fluent, was used to predict the heat transfer coefficient distributions on the holed surface, where there was no optical access during these tests. This also allowed investigation of additional rib configurations, and comparison of the pressure penalty associated with each design. The study showed that the turbulator configuration used allows the position and maximum level of heat transfer coefficient enhancement to be chosen by the engine designer. For the configurations tested heat transfer coefficient enhancement of up to 32% and 51% could be achieved on the holed surface and the ribbed surface respectively. For minimum additional pressure drop 45° ribs should be used.

Commentary by Dr. Valentin Fuster
2006;():1041-1053. doi:10.1115/GT2006-91245.

This paper presents detailed pressure measurements and discharge coefficient data for a long, low aspect ratio manifold; part of a novel blade cooling scheme. The cooling geometry, in which a series of racetrack passages are connected to a central plenum, provides high heat transfer coefficients in regions of the blade in good thermal contact with the outer blade surface. The Reynolds number changes along its length because of the ejection of fluid through a series of 19 transfer holes in a staggered arrangement, which are used to connect ceramic cores during the casting process. For rotation number RN = 0 the velocity down each hole remains almost constant. A correlation between hole discharge coefficient and Velocity Head Ratio is also presented. Pressure loss coefficients in the passage and through the holes are also discussed. A High Pressure (HP) rig was tested to investigate compressibility effects and expand the inlet Reynolds number range. A CFD model was validated against the experimental data, and then used to investigate the effects of rotation on the hole discharge coefficients. Results are presented for an inlet Reynolds number of 43477. At an engine representative rotation number of 0.08 corresponding buoyancy number of 0.17 there was little effect of rotation. However, at high rotational speeds secondary flows in the cooling passage and the exit plenum greatly reduce the hole discharge coefficient by increasing the local cross flow at the hole entrances and capping the hole exits in a manner similar to that seen in leading edge film-cooling geometries.

Commentary by Dr. Valentin Fuster
2006;():1055-1063. doi:10.1115/GT2006-91246.

Turbine blade coupons with three different surface treatments were exposed to deposition conditions in an accelerated deposition facility. The facility simulates the flow conditions at the inlet to a first stage high pressure turbine (T = 1150°C, M = 0.31). The combustor exit flow is seeded with dust particulate that would typically be ingested by a large utility power plant. The three coupon surface treatments included: (1) bare polished metal, (2) polished thermal barrier coating with bondcoat, and (3) unpolished oxidation resistant bondcoat. Each coupon was subjected to four successive 2 hour deposition tests. The particulate loading was scaled to simulate 0.02 ppmw (parts per million weight) of particulate over three months of continuous gas turbine operation for each 2 hour laboratory simulation (for a cumulative one year of operation). Three-dimensional maps of the deposit-roughened surfaces were created between each test, representing a total of four measurements evenly spaced through the lifecycle of a turbine blade surface. From these measurements the surface topology and roughness statistics were determined. Despite the different surface treatments, all three surfaces exhibited similar non-monotonic changes in roughness with repeated exposure. In each case, an initial build-up of isolated roughness peaks was followed by a period when valleys between peaks were filled with subsequent deposition. This trend is well documented using the average forward facing roughness angle in combination with the average roughness height as characteristic roughness metrics. Deposition-related mechanisms leading to spallation of the thermal barrier coated coupons are identified and documented.

Commentary by Dr. Valentin Fuster
2006;():1065-1074. doi:10.1115/GT2006-91250.

The cooling of rotor shrouds in the first stage of a high-pressure turbine requires special attention as flatter turbine inlet temperature profiles and more highly loaded blades result in increased thermal and mechanical stresses. The use of film cooling and/or internal convective cooling makes the rotor shroud heavier and oversized, restricting the maximum rotational speed. Alternative methods are therefore sought to achieve improved cooling of the shroud. This paper discusses the low speed experimental investigation of two ‘passive’ cooling concepts known as ‘rail cooling’ and ‘platform cooling’. It has been shown experimentally that the modified cooling method, namely the platform cooling, substantially improves the rotor shroud coolant distribution in the critical areas whilst employing significantly lower amounts of coolant.

Topics: Cooling , Turbines
Commentary by Dr. Valentin Fuster
2006;():1075-1082. doi:10.1115/GT2006-91257.

A TBC-coated turbine blade coupon was exposed to successive deposition in an accelerated deposition facility simulating flow conditions at the inlet to a first stage high pressure turbine (T = 1150°C, M = 0.31). The combustor exit flow was seeded with dust particulate that would typically be ingested by a large utility power plant. The turbine coupon was subjected to four successive 2 hour deposition tests. The particulate loading was scaled to simulate 0.02 ppmw (parts per million weight) of particulate over three months of continuous gas turbine operation for each 2 hour laboratory simulation (for a cumulative one year of operation). Three-dimensional maps of the deposit-roughened surfaces were created between each test, representing a total of four measurements evenly spaced through the lifecycle of a turbine blade surface. From these measurements, scaled models were produced for testing in a low-speed wind tunnel with a turbulent, zero pressure gradient boundary layer at Re = 750,000. The average surface heat transfer coefficient was measured using a transient surface temperature measurement technique. Stanton number increases initially with deposition but then levels off as the surface becomes less peaked. Subsequent deposition exposure then produces a second increase in St. Surface maps of St highlight the local influence of deposit peaks with regard to heat transfer.

Commentary by Dr. Valentin Fuster
2006;():1083-1094. doi:10.1115/GT2006-91273.

Pulsed film cooling was studied experimentally to determine its effect on film cooling effectiveness. The film cooling jets were pulsed using solenoid valves in the supply air line. Cases with a single row of cylindrical film cooling holes inclined at 35 degrees to the surface of a flat plate were considered at blowing ratios of 0.25, 0.5, 1.0, and 1.5 for a variety of pulsing frequencies and duty cycles. Temperature measurements were made using an infrared camera, thermocouples, and cold wire anemometry. Hot wire anemometry was used for velocity measurements. The local film cooling effectiveness was calculated based on the measured temperatures and the results were compared to baseline cases with continuous blowing. Phase locked flow temperature fields were determined from cold wire surveys. Pulsing at high frequencies helped to improve film cooling effectiveness in some cases by reducing overall jet liftoff. At lower frequencies, pulsing tended to have the opposite effect. With the present geometry and a steady mainflow, pulsing did not provide an overall benefit. The highest overall effectiveness was achieved with continuous jets and a blowing ratio of 0.5. The present results may prove useful for understanding film cooling behavior in engines, where mainflow unsteadiness causes film cooling jet pulsation.

Topics: Temperature , Cooling
Commentary by Dr. Valentin Fuster
2006;():1095-1104. doi:10.1115/GT2006-91274.

Pulsed film cooling was studied experimentally to determine its effect on film cooling effectiveness and heat transfer. The film cooling jets were pulsed using solenoid valves in the supply air line. Cases with a single row of cylindrical film cooling holes inclined at 35 degrees to the surface of a flat plate were considered at blowing ratios of 0.25, 0.5, 1.0, and 1.5 for a variety of pulsing frequencies and duty cycles. Temperature measurements were made using an infrared camera and thermocouples. The plate was equipped with constant flux surface heaters, and data were acquired for each flow condition with the plate both heated and unheated. The local film cooling effectiveness, Stanton numbers, and heat flux ratios were calculated and compared to baseline cases with continuous blowing and no blowing. Stanton number signatures on the surface provided evidence of flow structures including horseshoe vortices wrapping around the film cooling jets and vortices within the jets. Pulsing tends to increase Stanton numbers, and the effect tends to increase with pulsing frequency and duty cycle. Some exceptions were observed, however, at the highest frequencies tested. Overall heat flux ratios also show that pulsing tends to have a detrimental effect with some exceptions at the highest frequencies. The best overall film cooling was achieved with continuous jets and a blowing ratio of 0.5. The present results may prove useful for understanding film cooling behavior in engines, where mainflow unsteadiness causes film cooling jet pulsation.

Commentary by Dr. Valentin Fuster
2006;():1105-1118. doi:10.1115/GT2006-91318.

Numerical predictions and measurements of the flow structure and the Nusselt number behavior are presented for the endwall region of a linear blade cascade with contoured leading edge fillets. Two types of fillet profiles are employed: (a) a linear profile and (b) a circular profile. The fillets are placed at the junction of the blade leading edge and bottom endwall. The results with the fillets are also compared with the baseline case (predicted and measured) without any leading edge fillet. The pitchwise pressure gradient and shear stress on the endwall region are reduced for the filleted blades compared to the baseline case where there is no fillet. The flow turning from the axial direction (yaw angle) and the pitchwise flow velocity at the endwall boundary layer are also reduced in the vicinity of the fillets. In addition, the fillets weaken the leading edge horseshoe vortex by significantly reducing its vorticity and turbulent kinetic energy. Because of such changes in the endwall boundary layer, especially in the region where the fillets are located, the streamwise vorticity, turbulence kinetic energy, and total pressure losses in the passage vortex and the suction side leg vortex are reduced compared to the baseline. Endwall Nusselt numbers for the filleted cases are reduced upstream of the throat region.

Commentary by Dr. Valentin Fuster

Transition

2006;():1119-1126. doi:10.1115/GT2006-90044.

A transition model for describing wake-induced transition is presented based on the SST turbulence model by Menter and two dynamic equations for intermittency: one for near-wall intermittency and one for free-stream intermittency. In the Navier-Stokes equations, the total intermittency factor, which is the sum of the two, multiplies the turbulent viscosity computed by the turbulence model. The quality of the transition model is illustrated on the T106A test cascade for different levels of inlet free-stream turbulence intensity. The unsteady results are presented in space-time diagrams of shape factor, wall shear stress, momentum thickness and intermittency on the suction side. Results show the capability of the model to capture the physics of unsteady transition. Inevitable shortcomings are also revealed.

Commentary by Dr. Valentin Fuster
2006;():1127-1138. doi:10.1115/GT2006-90273.

This paper presents a thorough assessment for two of the contemporary CFD programs available for modeling and predicting nonfilm-cooled surface heat transfer distributions on turbine airfoil surfaces. The CFD programs are capable of predicting laminar-turbulent transition and have been evaluated and validated against five test cases with experimental data. The suite of test cases considered for this study consists of two flat plat cases at zero and non-zero pressure gradient and three linear-turbine-cascade test cases that are representative of modern high pressure turbine designs. The flat plate test cases are the ERCOFTAC T3A and T3C2, while the linear turbine cascade cases are the MARKII, the Virginia Polytechnic Institute (VPI), and the Von Karman Institute (VKI) turbine cascades. The numerical tools assessed in this study are 3D viscous Reynolds Averaged-Navier-Stokes (RANS) equations programs that employ a variety of one-equation and two-equation models for turbulence closure. The assessment study focuses on the one-equation Spalart and Allmaras and the two-equation shear stress transport K-ω turbulence models with the ability of modeling and predicting laminar-turbulent transition. The RANS 3D viscous codes are Numeca’s Fine Turbo and ANSYS-CFX’ CFX5. Numerical results for skin friction, surface temperature distribution and heat transfer coefficient from the CFD programs are compared to measured experimental data. Sensitivity of the predictions to free stream turbulence and to inlet turbulence boundary conditions is also presented. The results of the study clearly illustrate the superiority of using the laminar-turbulent transition prediction in improving the accuracy of predicting the heat transfer coefficient on the surfaces of high pressure turbine airfoils.

Commentary by Dr. Valentin Fuster
2006;():1139-1151. doi:10.1115/GT2006-90316.

A new model for predicting heat transfer in the transitional boundary layer of rough turbine airfoils is presented. The new model makes use of extensive experimental work recently published by the current authors. For the computation of the turbulent boundary layer a discrete element roughness model is combined with a two-layer model of turbulence. The transition region is modeled using an intermittency equation that blends between the laminar and turbulent boundary layer. Several intermittency functions are evaluated in respect of their applicability to rough-wall transition. To predict the onset of transition a new correlation is presented, accounting for the influence of free-stream turbulence and surface roughness. Finally the new model is tested against transitional rough-wall boundary layer flows on high-pressure and low-pressure turbine airfoils.

Commentary by Dr. Valentin Fuster
2006;():1153-1161. doi:10.1115/GT2006-90365.

The accurate numerical simulation of the flow through turbomachinery depends on the reliable prediction of laminar to turbulent boundary layer transition phenomena. The aim of this paper is to study the ability of the turbulent potential model to predict those non-equilibrium turbulent flows for several test cases. Within this model turbulent quantities are described by the turbulent scalar and turbulent vector potentials of the turbulent body force — the divergence of the Reynolds stress tensor. For model validation firstly flat plate test cases with different inlet turbulence intensities, zero pressure gradient and non-uniform pressure gradient distributions along the plate were calculated and compared by means of skin friction values measured in the experiments. Finally the model was validated by heat transfer measurement data obtained from a highly loaded transonic turbine guide vane cascade for different operating conditions.

Commentary by Dr. Valentin Fuster
2006;():1163-1172. doi:10.1115/GT2006-90403.

The experimental program was performed in the U.S. Air Force Academy water tunnel to visualize details of the unsteady flow structure in front of, within and downstream of a double-row array of shallow (h/D = 0.1) spherical and cylindrical dimples placed on a flat plate. The dimple projected diameter was 50.8 mm. The center of the first array was located at 88,0 mm downstream of the elliptically-shaped leading edge of the plate, the spanwise dimple pitch is 76.2 mm (Sz /D = 1.50). The second array was arranged in a staggered mode with the same spanwise pitch and the axial pitch between rows of 88.0 mm (Sx /D = 1.73). The diameter-based Reynolds number ReD ranged from 3,260 to 23,450, while the length-based Reynolds number Rex in front of the first array varied from 4,010 to 28,840. The laminar flow existed upstream of the first array for all flow conditions studied. In front of the second array the flow structure appeared to have either laminar or turbulent depending on the Reynolds number and streamline location. Flow visualizations were performed at 13 different flow speeds using the dye visualization technique. All recordings were made with a SONY-DCR VX2000 video camera. Adobe Premiere 6.5 software was used to analyze the flow characteristics using the slow motion feature. The results presented include details of the unsteady flow structure, bulk flow fluctuations and laminar-turbulent flow transition. The upstream vortex structures reduce bulk flow fluctuations beyond the second array making them smaller than those after the first array. At ReD > 13,000 the spherical dimples in the second array produce more significant bulk flow fluctuations than cylindrical dimples.

Topics: Flow (Dynamics)
Commentary by Dr. Valentin Fuster
2006;():1173-1179. doi:10.1115/GT2006-90455.

Boundary layers on concave surfaces differ from those on flat plates due to the presence of Taylor Goertler (T-G) vortices. These vortices cause momentum transfer normal to the blade’s surface and hence result in a more rapid development of the laminar boundary layer and a fuller profile than is typical of a flat plate. Transition of boundary layers on concave surfaces also occurs at a lower Rex than on a flat plate. Concave surface transition correlations have been formulated previously from experimental data, but they are not comprehensive and are based on relatively sparse data. The purpose of the current work was to attempt to model the physics of both the laminar boundary layer development and transition process in order to produce a transition model suitable for concave surface boundary layers. The development of the laminar boundary layer on a concave surface was modeled by considering the profiles at the upwash and downwash locations separately. The profiles of the boundary layers at these two locations were then combined to successfully approximate the spanwise averaged profile. The ratio of the boundary layer thicknesses at the two locations was found to be as great as 50 and this leads to laminar boundary layer shape factors as low as 1.3 and skin friction coefficients up to 12 times the value for a flat plate laminar boundary layer. Boundary layers therefore grow much more rapidly on concave surfaces than on flat plates. The transition model assumed that transition commenced in the upwash location boundary layer at the same transition inception Reθ observed on a flat plate. Transition at the downwash location then results from the growth of turbulent spots from the upwash location rather than through the initiation of spots. The model showed that initially curvature promotes transition because of the thickened upwash boundary layer, but for strong curvature the T-G vortices effectively stabilise the boundary layer and transition then occurs at a higher Reθ than on a flat plate. Results from the transition model were in broad agreement with experimental observations. The current work therefore provides a basis for the modeling of transition on concave surfaces.

Commentary by Dr. Valentin Fuster
2006;():1181-1190. doi:10.1115/GT2006-90543.

With this paper, results of a numerical investigation of the influence of the inlet condition variation on a stator vane suction side boundary layer and its separation tendencies are presented. The profile used for the examination is a so called high-lift-profile and designed for a laminar-turbulent transition over a steady separation bubble in a 1.5-stage low pressure turbine. Hence, the turbulence model must be capable for these effects. Especially, the stream line curvature has to be kept properly which leads to higher level turbulence models. The calculations were conducted with a two-dimensional Navier-Stokes solver using a finite volume discretisation scheme. The turbulence models used are the v′2 -f and the LCL turbulence model which are both of higher order. In the first part of the paper, wake free averaged inflow conditions were used. Through this, the influence of the mean flow on the bubble could be examined.

Commentary by Dr. Valentin Fuster
2006;():1191-1199. doi:10.1115/GT2006-90670.

We study the control of two-dimensional laminar separation bubbles on a flat plate at low Reynolds numbers, using two-dimensional DNS. A range of steady separation bubbles is obtained varying the pressure gradient. They are forced by a zero-mass flow, oscillatory wall blowing with different perturbation amplitudes and frequencies. The reduction in bubble length as a function of frequency has two minima for sufficient high amplitudes. One of them is related to the Kelvin-Helmholtz instability of the separated boundary layer, while the other, most effective one, is here denoted as the low-frequency regime. In this regime large vortices are created which are not a consequence of an instability of the original bubble. On the contrary the forcing creates an unsteady separation bubble which evolves into a large vortex. These vortices have large radii and attach to the wall due to their self-induced pressure field while convecting across the adverse pressure gradient zone. Scaling relations for the effect of the forcing are proposed and tested.

Commentary by Dr. Valentin Fuster
2006;():1201-1210. doi:10.1115/GT2006-90754.

Flow measurements were made on a highly loaded low pressure turbine blade in a low-speed linear cascade facility. The blade has a design Zweifel coefficient of 1.34 with a peak pressure coefficient near 47% axial chord (mid-loaded). Flow and surface pressure data were taken for Rec = 20,000 with 3% inlet freestream turbulence. For these operating conditions, a large separation bubble forms over the downstream portion of the blade suction surface, extending from 59% to 86% axial chord. A Single-element hotfilm measurements were acquired to clearly identify the role of boundary layer transition in this separated region. Higher-order turbulence statistics were used to identify transition and separation zones. Similar measurements were also made in the presence of unsteady forcing using pulsed vortex generator jets just upstream of the separation bubble (50% cx ). Measurements provide a comprehensive picture of the interaction of boundary layer transition and separation in this unsteady environment. Similarities between pulsed flow control and unsteady wake motion are highlighted.

Commentary by Dr. Valentin Fuster
2006;():1211-1220. doi:10.1115/GT2006-90809.

The continuous tendency in modern aeroengine gas turbines towards reduction of blade count and ducts length may lead to aerodynamic loading increase beyond the limit of boundary layer separation. For this reason boundary layer separation control methods, up to now mostly employed in external aerodynamics, begin to be experimented in internal flows applications. The present paper reports the results of a detailed experimental study on low profile vortex generators used to control boundary layer separation on a large-scale flat plate with prescribed adverse pressure gradients. Inlet turbulent boundary layer conditions and pressure gradients are representative of aggressive turbine intermediate ducts. This activity is part of a joint European research program on Aggressive Intermediate Duct Aerodynamics (AIDA). The pressure gradients on the flat plate are generated by increasing the aperture angle of a movable wall opposite to the flat plate. To avoid separation on the movable wall, boundary layer suction is applied on it. Complementary measurements (surface static pressure distributions, surface flow visualizations by means of wall mounted tufts, instantaneous and time-averaged velocity fields in the meridional and cross-stream planes by means of Particle Image Velocimetry) have been used to survey the flow with and without vortex generators. Three different pressure gradients, which induce turbulent separation in absence of boundary layer control, were tested. Vortex generators height and location effects on separation reduction and pressure recovery increase were investigated. For the most effective VGs configurations detailed analyses of the flow field were performed, that demonstrate the effectiveness of this passive control device to control separation in diffusing ducts. Particle Image Velocimetry vector and vorticity plots illustrate the mechanisms by which the vortex generators transfer momentum towards the surface, re-energizing the near-wall flow and preserving the boundary layer from separation.

Commentary by Dr. Valentin Fuster
2006;():1221-1230. doi:10.1115/GT2006-91018.

In this paper, transition in a separation bubble is examined through numerical simulation. The flow Reynolds number and streamwise pressure distribution are typical of the conditions encountered on the suction side of low-pressure turbine blades of gas-turbine engines. The spatial and temporal resolutions utilized in the present computations correspond to a coarse direct numerical simulation, wherein the majority of turbulence scales, including the inertial subrange, are adequately resolved. The accuracy of the simulation results is demonstrated through favorable comparisons with experimental data corresponding to the same flow conditions. The results of the simulation show linear Tollmien-Schlichting (T-S) instability growth downstream of the point of separation, leading to the roll-up of spanwise vorticity into disctete vortical structures, characteristic of Kelvin-Helmholtz (K-H) instability growth. The extent of cross-stream momentum exchange associated with packets of amplified T-S waves is examined, along with details of the time-periodic breakdown into turbulence occurring upon the development of the K-H instability. Reynolds-averaged properties of the separation bubble are presented, and provide evidence of the strong three-dimensional nature of the reattachment process.

Commentary by Dr. Valentin Fuster
2006;():1231-1242. doi:10.1115/GT2006-91044.

An experimental study was performed to examine the effect of a dielectric barrier discharge plasma flow control actuator. The actuator was applied to a flat plate boundary layer in a pressure distribution approximating that of the suction side of an aft-loaded low-pressure turbine blade. Particle image velocimetry was used to obtain two-dimensional velocity field data on four boundary layers with velocity profiles that included fully attached laminar, turbulent, nominally separated and high-separated. Values of Rec ranged from 4500 to 75,000. For the separated boundary-layers the actuator was located downstream of the separation point. Four E-fields were characterized for each case. The actuator was able to induce boundary layer reattachment even when located downstream of the initial separation point. The length scale over which this re-attachment occurs was strongly dependent upon the applied E-field. The actuators also caused significant reduction in the overall boundary layer thickness as a result of the modification of the near-wall static pressure distribution.

Commentary by Dr. Valentin Fuster
2006;():1243-1250. doi:10.1115/GT2006-91103.

This study investigates wake-induced bypass transition of a flat-plate boundary layer that experiences favorable and adverse pressure gradients. The effect of free-stream turbulence is also examined. This paper mainly focuses on how the pressure gradients affect the transitional behavior of the wake-disturbed boundary layer with and without the influence of free-stream turbulence. A wall-contouring of the test duct generates the flow acceleration and deceleration on the flat plate, emulating aft-loading or front-loading pressure distributions of actual turbine blades. A spoked-wheel type wake generator is used to produce wakes moving over the boundary layer. Detailed boundary layer measurements are performed by use of a hot-wire anemometer. In addition to velocity fluctuations, which clearly indicate transition process, intermittency factors are obtained and compared with the results given by a transition model. Noticeable differences in transitional behavior are observed between the cases with and without the enhanced free-stream turbulence. It is also confirmed that the wake width as well as the direction of the bar movement are influential factors to the bypass transition.

Commentary by Dr. Valentin Fuster
2006;():1251-1259. doi:10.1115/GT2006-91166.

Accurate transition onset modeling is a fundamental part of modern turbomachinery designs, where bypass transition is the dominant mechanism of transition to turbulence. Despite this situation a range of transition onset models exist primarily based upon both integral and local parameters within the boundary layer. All such transition models have empirical origins. To date the relationships between such models has not been forthcoming and hence lack of physical understanding of the transition process is evident. This paper details a new approach to transition modeling and provides a theoretically based approach to transition onset prediction by invoking a single principle developed within constructal theory. We not only present a new model but also demonstrate the equivalence between existing models by implementing the same theory. Such understanding of the transition onset problem may provide a new perspective towards more theoretically based transition onset models rather than empirical ones, although much work remains to be done in understanding the receptivity mechanisms within a laminar boundary layer.

Topics: Boundary layers
Commentary by Dr. Valentin Fuster
2006;():1261-1269. doi:10.1115/GT2006-91193.

Laminar separation was investigated experimentally on a flat plate under a strongly diffusing self-similar pressure distribution. This gave a long and thin laminar separation bubble. Boundary layer velocity traverses were performed at numerous longitudinal stations. Using a single hot wire a combination of individual traces, phase averaging and time averaging was used. To supplement this, an array of microphones was installed to give instantaneous contours of pressure perturbation and to investigate the time dependent flow features. Microphone data were consistent with the strong amplification, under the adverse pressure gradient, of instabilities predicted far upstream of the separation point. Driven at the Tollmien-Schlichting (T-S) frequency, these instabilities grew into turbulent spots developing in the shear layer of the separation bubble. Reattachment of the bubble was caused by transition of the separated shear layer. The waves were strongest in the later stages of transition. Once the coherence was lost, in a turbulent layer, the amplitude became diminished. Wake disturbances were injected into the flow and traced through the flow field. The wake interaction resulted in turbulent patches which penetrated to the wall. Following the patches was the calmed region, detectable as a region of reduced wave activity in the transition region following each turbulent strip. For a short time at the end of the calmed region the viscous instability waves continued to propagate for a considerable distance downstream, in concert with the calmed region, in an otherwise turbulent zone.

Commentary by Dr. Valentin Fuster

Internal Air and Seals

2006;():1271-1282. doi:10.1115/GT2006-90033.

Thermal and flow field in a rotor-stator system around a single stage HP-turbine disk are investigated. In the previous work, the authors’ group applied a newly developed conjugate analysis method to the system at a steady-state condition, compared the thermal field with the measured data, and concluded that a satisfactory level of accuracy can be achieved with the proposed method at the steady-state condition. The present paper focuses on a heat and fluid flow in the same rotor-stator system, but at a transient condition which simulates a typical accelerating and decelerating schedule expected in actual aero-engine operations. First, detailed measurement of secondary-air and metal temperatures around the turbine disk is carried out in the transient condition. Then, the 3D fluid / solid conjugate analysis is applied to the whole accelerating and decelerating schedule. The results show that the transient thermal behavior of the rotor-stator system is well captured and reproduced with the present conjugate method.

Commentary by Dr. Valentin Fuster
2006;():1283-1290. doi:10.1115/GT2006-90035.

Oil containment is a critical design requirement that affects overall system safety and reliability of gas turbine engines. This paper examines a new method to enhance oil containment by use of an improved de-oiler that creates a favorable bearing compartment differential pressure environment even at low power settings. Typically gas turbine engines require seals to contain oil within the bearing compartment. These seals, both contacting and non-contacting configuration styles, rely on secondary airflow to buffer the sealing interface and force oil mist and droplets back into the compartment. This is not difficult to achieve at high or moderate power conditions since there is generally sufficient air flow and pressure available to meet the sealing requirements. However, at idle conditions, the engine low-pressure compressor (LPC) may not turn fast enough to produce sufficient airflow to buffer the seals. To address these concerns the authors propose a method where the de-oiler creates a vacuum at idle speed, which results in favorable compartment seal differential pressures and also acts as a restrictor at higher speeds, where limiting the contact pressure and increasing the service life of mechanical seals become desirable design goals. The paper will examine a specific case study with both analytical and experimental results.

Topics: Containment
Commentary by Dr. Valentin Fuster
2006;():1291-1300. doi:10.1115/GT2006-90132.

This paper compares heat transfer measurements from a pre-swirl rotor-stator experiment with 3D steady state results from a commercial CFD code. The measured distribution of Nusselt number on the rotor surface was obtained from a scaled model of a gas turbine rotor-stator system, where the flow structure is representative of that found in an engine. Computations were carried out using a coupled multigrid RANS solver with a high-Reynolds-number k-ε/k-ω turbulence model. Previous work has identified three parameters governing heat transfer: rotational Reynolds number (Reφ ), pre-swirl ratio (βp ) and the turbulent flow parameter (λT ). For this study rotational Reynolds numbers are in the range 0.8×106 < Reφ < 1.2×106 . The turbulent flow parameter and pre-swirl ratios varied between 0.12 < λT < 0.38 and 0.5 < βp < 1.5, which are comparable to values that occur in industrial gas turbines. At high coolant flow rates, computations have predicted peaks in heat transfer at the radius of the pre-swirl nozzles. These were discovered during earlier experiments and are associated with the impingement of the pre-swirl flow on the rotor disc. At lower flow rates, the heat transfer is controlled by boundary-layer effects. The Nusselt number on the rotating disc increases as either Reφ or λT increases, and is axisymmetric except in the region of the receiver holes, where significant two-dimensional variations are observed. The computed velocity field is used to explain the heat transfer distributions observed in the experiments. The regions of peak heat transfer around the receiver holes are a consequence of the route taken by the flow. Two routes have been identified: “direct”, whereby flow forms a stream-tube between the inlet and outlet; and “indirect”, whereby flow mixes with the rotating core of fluid. Two performance parameters have been calculated: the adiabatic effectiveness for the system, Θb,ab , and the discharge coefficient for the receiver holes, CD . The computations show that, although Θb,ab increases monotonically as βp increases, there is a critical value of βp at which CD is a maximum.

Commentary by Dr. Valentin Fuster
2006;():1301-1311. doi:10.1115/GT2006-90143.

The uncertainty in heat transfer data derived from theoretical turbomachinery disc temperatures has been investigated using Monte Carlo simulation methods. Two specific test cases of relevance to turbine disc heat transfer have been considered, a free disc with laminar flow and a free disc with turbulent flow. This study explores and quantifies uncertainties for the test cases considered by looking at three factors: measurement noise, the order of magnitude of the polynomial fit used for interpolation through the data points and the number of data points utilised within the polynomial. The local heat flux uncertainty arising from using a 5th order polynomial for the laminar case varied between approximately ± 40 % and ± 600 % and for the turbulent case between approximately ± 4 % and ± 1000 %. This study provides guidelines for similar analyses and the resulting uncertainty and how test data can be presented.

Commentary by Dr. Valentin Fuster