bluuebunny/crossref_metadata_2025 huggingface dataset:
Leading-Edge Film-Cooling Physics: Part II — Heat Transfer Coefficient
Authors: York William D. , Leylek James H.
Published: February 2009

A comprehensive study of film cooling on a turbine airfoil leading edge was performed with a documented, well-tested computational methodology. In this paper, numerically predicted heat transfer coefficients on the film-cooled leading edge are compared with experimental data from the open literature. The results are presented as the ratio of heat transfer coefficient with film cooling to that without film cooling, and the physics behind the surface results are discussed. The leading edge model was a half-cylinder in shape with a bluff afterbody to match the validation experiment, and other geometric parameters matched those of Part I of this study. Coolant at a density equal to that of the mainstream flow was injected through three rows of cylindrical film-cooling holes. One row of holes was centered on the stagnation line of the cylinder, and the other two rows were located ±3.5 hole diameters off stagnation. The downstream rows were staggered such that they were centered laterally between holes in the stagnation row. The holes were inclined at 20° with the surface, and made a 90° angle with the streamwise direction (radial injection). Four average blowing ratios were simulated in the range of 0.75 to 1.9, corresponding to the same momentum flux ratios as in Part I of this work. The multi-block, unstructured numerical grid was characterized by high quality and density, especially in the near wall region, in order to minimize error in predictions of the heat transfer. A fully-implicit scheme was used to solve the steady Reynolds-averaged Navier-Stokes equations, and a realizable k-ε model provided turbulence closure. A two-layer near-wall treatment allowed the resolution of the viscous sublayer for maximum accuracy in the prediction of the wall heat transfer coefficient. The numerical predictions exhibited generally good agreement with experimental data. The heat transfer coefficient was observed to increase sharply aft of the holes in the downstream rows. When coupled with the adiabatic effectiveness results of the first paper in this series, it is evident that a systematic computational methodology may be effectively applied to investigate and understand the complicated leading edge film-cooling problem.
Film Cooling Experiments With Flow Introduced Upstream of a First Stage Nozzle Guide Vane Through Slots of Various Geometries
Authors: Oke Rohit A. , Simon Terrence W.
Published: February 2009

This paper describes the advantages of introducing film cooling flow through the endwall upstream of the first stage nozzle guide vane. To perform these studies, a linear cascade is built. It consists of three vanes and two endwalls that form two passages. One endwall is flat and the other is contoured from upstream of the leading edge, continuing through the passage. The approach flow is of high turbulence and large length scale, representative of the engine combustor exit flow. Film cooling flow is introduced through two successive rows of slots, a single row of slots and slots that have particular area distributions in the pitchwise direction. Measurements are taken by heating the film cooling flow slightly above the main flow temperature and recording temperature distributions in the film cooling flow-main flow mixing zone at various axial planes. The single and double slot injection cases represent base-line injection geometries. They show that at lower ratios of coolant to main flow momentum fluxes, film cooling flow migrates toward the suction side due to secondary flow. At higher ratios, the pressure side endwall region is cooled more effectively. Observations are drawn by comparing the baseline injection cases with cases of different geometries for which slots are blocked partially to re-distribute mass and momentum injection rates of the emerging flow. The downstream evolutions of temperature contours are discussed. The idea is to utilize secondary flows to control pitchwise coolant distributions.
Free-Stream Turbulence Effects on Film Cooling With Shaped Holes
Authors: Saumweber Christian , Schulz Achmed , Wittig Sigmar
Published: February 2009

A comprehensive set of generic experiments has been conducted to investigate the effect of elevated free-stream turbulence on film cooling performance of shaped holes. A row of three cylindrical holes as a reference case, and two rows of holes with expanded exits, a fanshaped (expanded in lateral direction), and a laidback fanshaped hole (expanded in lateral and streamwise direction) have been employed. With an external (hot gas) Mach number of Mam = 0.3 operating conditions are varied in terms of free-stream turbulence intensity (up to 11%), integral length scale at constant turbulence intensity (up to 3.5 hole inlet diameters), and blowing ratio. The temperature ratio is fixed at 0.59 leading to an engine-like density ratio of 1.7. The results indicate that shaped and cylindrical holes exhibit very different reactions to elevated free-stream turbulence levels. For cylindrical holes film cooling effectiveness is reduced with increased turbulence level at low blowing ratios whereas a small gain in effectiveness can be observed at high blowing ratios. For shaped holes, increased turbulence intensity is detrimental even for the largest blowing ratio (M = 2.5). In comparison to the impact of turbulence intensity the effect of varying the integral length scale is found to be of minor importance. Finally the effect of elevated free-stream turbulence in terms of heat transfer coefficients was found to be much more pronounced for the shaped holes.
Numerical Investigation of Internal Crossflow Film Cooling
Authors: Adami P. , Martelli F. , Montomoli F. , Saumweber C.
Published: February 2009

A full-3D unstructured solver is applied for the investigation of the physics involved in the flow of modern film cooling devices. The numerical simulation is based on a TVD upwind finite volume method that exploits the implicit time-marching approach. A conventional two equations eddy viscosity closure is considered for the turbulent flow field without wall-functions. The present application aims to investigate and discuss the flow field physics as obtained from a numerical RANS (Reynolds Averaged Navier-Stokes) simulation comparing different cooling duct systems devices. The CFD outcome is discussed against experiments considering the discharge coefficient as a means to quantify the accuracy of the numerical simulation. Particular attention is focused on the geometrical discretization, on the grid characteristics and on the capabilities of CFD for an efficient and realistic modelling of the flow field. The basic features of the film cooling technique are addressed considering the experimental test configurations investigated by Karlsruhe University for cylindrical, fanshaped and laid back fan shaped configurations with cross/parallel flows arrangement.
The Effect of Freestream Turbulence on Film Cooling Adiabatic Effectiveness
Authors: Mayhew James E. , Baughn James W. , Byerley Aaron R.
Published: February 2009

The film-cooling performance of a flat plate in the presence of low and high freestream turbulence is investigated using liquid crystal thermography. High-resolution distributions of the adiabatic effectiveness are determined over the film-cooled surface of the flat plate using the hue method and image processing. Three blowing rates are investigated for a model with three straight holes spaced three diameters apart, with density ratio near unity. High freestream turbulence is shown to increase the area-averaged effectiveness at high blowing rates, but decrease it at low blowing rates. At low blowing ratio, freestream turbulence clearly reduces the coverage area of the cooling air due to increased mixing with the main flow. However, at high blowing ratio, when much of the jet has lifted off in the low turbulence case, high freestream turbulence turns its increased mixing into an asset, entraining some of the coolant that penetrates into the main flow and mixing it with the air near the surface.
Evaluation of Pressure Side Film Cooling With Flow and Thermal Field Measurements: Part II — Turbulence Effects
Authors: Cutbirth J. Michael , Bogard David G.
Published: February 2009

This study focused on the film cooling performance on the pressure side of a turbine vane subjected to high mainstream turbulence levels, with and without showerhead blowing. Whereas previous studies have measured the adiabatic effectiveness and heat transfer at the surface of the airfoil, the goal of this study was to examine the flow and thermal fields above the surface. These measurements included flow visualization, thermal profiles, and laser Doppler velocimetry. For comparison, adiabatic effectiveness was also measured. A mainstream turbulence level of Tu∞ = 20%, with integral length scale of seven hole diameters, was used. Particularly insightful is the discovery that the large scale high mainstream turbulence causes a lateral oscillation of coolant jet resulting in a much wider time average distribution of coolant. Even with high mainstream turbulence, showerhead blowing was found to still cause a significantly increased dispersion of the pressure side coolant jets.
Influence of Discrete Pin Shaped Surface Roughness (P-Pins) on Heat Transfer and Aerodynamics of Film Cooled Aerofoil
Authors: Guo S. M. , Oldfield M. L. G. , Rawlinson A. J.
Published: February 2009

The influence of localized pin-shaped surface roughness (P-Pins) on heat transfer and aerodynamics of a fully film cooled engine aerofoil has been studied in a transonic annular cascade. The “P-Pins”, present on some casting film cooled turbine blades and vanes, are the residues left in the manufacturing process. This paper investigates the effect of the P-Pins on the aerodynamic performance and measures the heat transfer consequences both for the aerofoils and the P-Pins. The effect on performance was determined independently on the pressure and suction surface of the aerofoil. For comparison, the aerofoil without P-Pins was also tested to provide baseline results. The measurements have been made at engine representative Mach and Reynolds numbers. Wide band liquid crystal and direct heat flux gauge technique were employed in the heat transfer tests. A four-hole pyramid probe was used to obtain the aerodynamic data. The aerodynamic and thermodynamic characteristics of the coolant flow have been modelled to represent engine conditions by using a heavy “foreign gas” (30.2% SF6 and 69.8% Ar by weight). The paper concludes that P-Pins as usually placed on the blade do not have detrimental effects to the heat transfer performance of film-cooled aerofoil. P-Pins, located in thick boundary layer regions of the aerofoil, such as the later portion of the suction surface, do not cause any reduction of aerofoil aerodynamic efficiency. For contrast, the P-Pins located in the thin boundary layer regions on the pressure side of the aerofoil cause noticeably more losses.
Correlation of Film Cooling Effectiveness From Thermographic Measurements at Engine Like Conditions
Authors: Baldauf S. , Scheurlen M. , Schulz A. , Wittig S.
Published: February 2009

Adiabatic film cooling effectiveness on a flat plate surface downstream of a row of cylindrical holes is investigated. Highly resolved two dimensional surface data were measured by means of infrared thermography and carefully corrected for local conduction and radiation effects [1]. These locally acquired data are laterally averaged to give the streamwise distributions of the effectiveness. An independent variation of the flow parameters blowing rate, density ratio, and turbulence intensity as well as the geometrical parameters streamwise ejection angle and hole spacing is examined. The influences of these parameters on the laterally effectiveness is discussed and interpreted with the help of surface distributions of effectiveness and heat transfer coefficients presented in earlier publications [1, 2]. Besides the known jet in cross-flow behavior of coolant ejected from discrete holes, these data demonstrate the effect of adjacent jet interaction and its impact on jet lift-off and adiabatic effectiveness. In utilizing this large matrix of measurements the effect of single parameters and their interactions are correlated. The important scaling parameters of the effectiveness are shaped out during the correlation process and are discussed. The resulting new correlation is designed to yield the quantitatively correct effectiveness as a result of the interplay of the jet in crossflow behavior and the adjacent jet interaction. It is built modularly to allow for future inclusion of additional parameters. The new correlation is valid without any exception within the full region of interest, reaching from the point of the ejection to far downstream, for all combinations of flow and geometry parameters.
Gas Turbine Engine Durability Impacts of High Fuel-Air Ratio Combustors: Part 2 — Near Wall Reaction Effects on Film-Cooled Heat Transfer
Authors: Kirk Daniel R. , Guenette Gerald R. , Lukachko Stephen P. , Waitz Ian A.
Published: February 2009

As commercial and military aircraft engines approach higher total temperatures and increasing overall fuel-to-air ratios, the potential for significant chemical reactions on a film-cooled surface is enhanced. Currently there is little basis for understanding the effects on aero-performance and durability due to such secondary reactions. A shock tube experiment was employed to generate short duration, high temperature (1000–2800 K) and pressure (6 atm.) flows over a film-cooled flat plate. The test plate contained two sets of 35° film cooling holes that could be supplied with different gases, one side using air and the other nitrogen. A mixture of ethylene and argon provided a fuel rich freestream that reacted with the air film resulting in near wall reactions. The relative increase in surface heat flux due to near wall reactions was investigated over a range of fuel levels, momentum blowing ratios (0.5–2.0), and Damko¨hler numbers (ratio of flow to chemical time scales) from near zero to 30. For high Damko¨hler numbers, reactions had sufficient time to occur and increased the surface heat flux by 30 percent over the inert cooling side. When these results are appropriately scaled, it is shown that in some situations of interest for gas turbine engine environments significant increases in surface heat flux can be produced due to chemical reactions in the film-cooling layer. It is also shown that the non-dimensional parameters Damko¨hler number (Da), blowing ratio (B), heat release potential (H*), and scaled heat flux (Qs) are the appropriate quantities to predict the augmentation in surface heat flux that arises due to secondary reactions.
Experimental and Numerical Study of Heat Transfer in a Gas Turbine Combustor Liner
Authors: Bailey Jeremy C. , Intile John , Fric Thomas F. , Tolpadi Anil K. , Nirmalan Nirm V. , Bunker Ronald S.
Published: February 2009

Experiments and numerical simulations were conducted to understand the heat transfer characteristics of a stationary gas turbine combustor liner cooled by impingement jets and cross flow between the liner and sleeve. Heat transfer was also aided by trip-strip turbulators on the outside of the liner and in the flowsleeve downstream of the jets. The study was aimed at enhancing heat transfer and prolonging the life of the combustor liner components. The combustor liner and flow sleeve were simulated using a flat plate rig. The geometry has been scaled from actual combustion geometry except for the curvature. The jet Reynolds number and the mass-velocity ratios between the jet and cross flow in the rig were matched with the corresponding combustor conditions. A steady state liquid crystal technique was used to measure spatially resolved heat transfer coefficients for the geometric and flow conditions mentioned above. The heat transfer was measured both in the impingement region as well as over the turbulators. A numerical model of the combustor test rig was created that included the impingement holes and the turbulators. Using CFD, the flow distribution within the flow sleeve and the heat transfer coefficients on the liner were both predicted. Calculations were made by varying the turbulence models, numerical schemes, and the geometrical mesh. The results obtained were compared to the experimental data and recommendations have been made with regard to the best modeling approach for such liner-flow sleeve configurations.
Augmentation of Stagnation Region Heat Transfer Due to Turbulence From an Advanced Dual-Annular Combustor
Authors: Van Fossen G. James , Bunker Ronald S.
Published: February 2009

Heat transfer measurements have been made in the stagnation region of a flat plate with an elliptical leading edge. The radius of curvature at the stagnation point was similar to that of a first stage turbine vane airfoil used in a large commercial high-bypass turbofan engine. The airfoil was mounted downstream of an arc segment of a dual-annular combustor similar to the type used in an advanced turbine engine. Testing was done in air at atmospheric temperature and at pressures up to 376 kPa to simulate the vane leading edge Reynolds number seen in the engine. Spanwise average stagnation region heat transfer was measured with an electrically heated aluminum strip. Turbulence intensity, length scale and isotropy were measured using standard 2-wire hot wire probes. The combustor contained two annular rows of fuel-air swirlers which were aligned in the radial direction. Both heat transfer and hot wire data were taken at two circumferential positions; one directly downstream of a pair of swirlers and one half way between two pairs of swirlers. Reynolds number based on vane leading edge diameter was varied from 51000 to 160000. The maximum Reynolds number for turbulence measurements was limited to 87000. Turbulence intensity averaged over all test conditions was found to be 31.6%. Average axial, integral length scale was 1.29 cm, which gave a length scale-to-leading edge diameter ratio of 1.08. The turbulence was found to be nearly isotropic with the average ratio of axial to circumferential fluctuating components of 1.15. Heat transfer augmentation above laminar levels was found to vary from 34 to almost 59% depending on the Reynolds number. No effect of circumferential position was found. The heat transfer augmentation was found to be well predicted by a correlation derived from grid generated turbulence.
Turbine Tip and Shroud Heat Transfer and Loading: Part A — Parameter Effects Including Reynolds Number, Pressure Ratio, and Gas to Metal Temperature Ratio
Authors: Polanka Marc D. , Hoying Donald A. , Meininger Matthew , MacArthur Charles D.
Published: February 2009

Turbine tip and shroud flow and heat transfer are some of the most complex, yet important, issues in turbine design. Most of the work performed to date has been performed in linear cascades and has investigated such items as the effect of tip geometries and turbulence on tip and shroud pressure and heat transfer. There have been very few full annulus or rotating measurements in the literature. Experimental measurements have been made on a single stage high pressure turbine at the US Air Force Turbine Research Facility (TRF) to aid in the understanding of this phenomena. The TRF is a full scale, rotating rig that operates at matched flow conditions to the true turbine environment. Heat flux measurements were acquired with both Pyrex insert strip and button gauges while the pressure measurements were taken with surface mounted Kulite® pressure transducers. This paper presents one of the first full rotating, simultaneous pressure and heat transfer measurements to be taken in the turbine tip shroud region. These measurements provide some of the details needed for accurately quantifying the true flow condition in this complex flow regime. Comparisons between the present data and the existing 2D cascade data were made. This investigation quantified the effects of Reynolds number, inlet temperature, turbine pressure ratio and inlet flow temperature profiles. This provides a benchmark data set for validation of numerical codes.
Effects of High Free-Stream Turbulence on the Near-Wall Flow and Heat/Mass Transfer on the Endwall of a Linear Turbine Rotor Cascade
Authors: Lee Sang Woo , Jun Sang Bae , Park Byung-Kyu , Lee Joon Sik
Published: February 2009

Experimental data are presented which describe the effects of a combustor-level high free-stream turbulence on the near-wall flow structure and heat/mass transfer in the endwall region of a linear high-turning turbine rotor cascade. The endwall flow structure is visualized by employing the partial- and total-coverage oil-film technique, and heat/mass transfer rate is measured by the naphthalene sublimation method. A turbulence generator is designed to provide a turbulent boundary layer flow which has free-stream turbulence intensity and integral length scale of 14.7% and 80 mm, respectively, at the entrance of the turbine cascade. The surface flow visualization shows that the high free-stream turbulence has little effect on the attachment line, but alters the separation line noticeably. Under high free-stream turbulence, the incoming near-wall flow upstream of the adjacent separation lines collides at a shallower angle with the suction surface. A weaker lift-up force arising from this more oblique collision results in the narrower suction-side corner vortex area in the high turbulence case. The high free-stream turbulence enhances the heat/mass transfer in the central area of the turbine passage, but only a slight augmentation is found in the endwall regions adjacent to the leading and trailing edges. Therefore, the high free-stream turbulence makes the endwall heat load more uniform. It is also observed that the heat/mass transfers along the locus of the pressure-side leg of the leading-edge horseshoe vortex and along the suction-side corner are influenced most strongly by the high free-stream turbulence. The endwall surface is classified into seven different regions based on the local heat/mass transfer distribution, and the effects of the high free-stream turbulence on the local heat/mass transfer in each region are discussed in detail.
Prediction and Measurement of the Flow and Heat Transfer Along the Endwall and Within an Inlet Vane Passage
Authors: Pasinato Hugo D. , Liu Zan , Roy Ramendra P. , Howe W. Jeffrey , Squires Kyle D.
Published: February 2009

Numerical simulations and laboratory measurements are performed to study the flow field and heat transfer in a linear cascade of turbine vanes. The vanes are scaled-up versions of a turbine engine inlet vane but simplified in that they are untwisted and follow the mid-span airfoil shape of the engine vane. The hub endwall is axially profiled while the tip endwall is flat. The hub endwall comprises the focus of the heat transfer investigation. Configurations are considered with and without air injection through three discrete angled (25 degrees to the main flow direction) slots upstream of each vane. The freestream turbulence intensity at the vane cascade inlet plane is 11 (± 2) percent, as measured by a single hot-wire placed perpendicular to the mean flow. The transient thermochromic liquid crystal technique is used to measure the convective heat transfer coefficient at the hub endwall for the baseline case (without air injection through the slots), and the heat transfer coefficient and cooling effectiveness at the same endwall for the cases with air injection at two blowing ratios. Miniature Kiel probes are used to measure the distribution of total pressure upstream of, within, and downstream of one vane passage. Numerical simulations are performed of the incompressible flow using unstructured grids. Hybrid meshes comprised of prisms near solid surfaces and tetrahedra away from the wall are used to resolve the solutions, with mesh refinement up to approximately 2 million cells. For all calculations, the first grid point is within one viscous unit of solid surfaces. A Boussinesq approximation is invoked to model the turbulent Reynolds stresses, with the turbulent eddy viscosity obtained from the Spalart-Allmaras one-equation model. The turbulent heat flux is modeled via Reynolds analogy and a constant turbulent Prandtl number of 0.9. The simulations show that endwall axial profiling results in flow reversal upstream of the vane, an effect that lowers the Stanton number for the baseline flow near the vane leading edge compared to our previous work in a flat-endwall geometry. Predictions of the total pressure loss coefficient show that the peak levels are higher than those measured.
Flow and Heat Transfer Predictions for a Flat-Tip Turbine Blade
Authors: Yang Huitao , Acharya Sumanta , Ekkad Srinath V. , Prakash Chander , Bunker Ron
Published: February 2009

Numerical calculations are performed to simulate the tip leakage flow and heat transfer on the GE-E3 High-Pressure-Turbine (HPT) rotor blade. The calculations are performed for a single blade with periodic conditions imposed along the two boundaries in the circumferential-pitch direction. Cases considered are a flat blade tip at three different tip gap clearances of 1%, 1.5% and 2.5% of the blade span. The numerical results are obtained for two different pressure ratios (ratio of inlet total pressure to exit static pressure) of 1.2 and 1.32 and an inlet turbulence level of 6.1%. To explore the effect of turbulence models on the heat transfer results, three different models of increasing complexity and computational effort (standard high Re k-ε model, RNG k-ε and Reynolds Stress Model) are investigated. The predicted tip heat transfer results are compared with the experimental data of Azad [1], and show satisfactory agreement with the data. Hear transfer predictions for all three turbulence models are comparable, and no significant improvements are obtained with the Reynolds-stress model.
Numerical Simulation of Flow and Heat Transfer Past a Turbine Blade With a Squealer-Tip
Authors: Yang Huitao , Acharya Sumanta , Ekkad Srinath V. , Prakash Chander , Bunker Ron
Published: February 2009

Numerical calculations are performed to simulate the tip leakage flow and heat transfer on the squealer (recessed) tip of GE-E3 turbine rotor blade. A squealer tip with a 3.77% recess of the blade span is considered in this study, and the results are compared with the predictions for a flat-tip blade. The calculations have been performed for an isothermal blade with an overall pressure ratio of 1.32, an inlet turbulence intensity of 6.1%, and for three different tip gap clearances of 1%, 1.5% and 2.5% of the blade span. These conditions correspond to the experiments reported by Azad et al. [1]. The calculations have been performed for three different turbulence models (the standard high Re k-ε model, the RNG k-ε and the Reynolds Stress Model) in order to assess the capability of the models in correctly predicting the blade heat transfer. The predictions show good agreement with the experimental data, with the Reynolds stress model calculations clearly providing the best results. Substantial reductions in the tip heat transfer and leakage flow is obtained with the squealer tip configuration. With the squealer tip, the heat transfer coefficients on the shroud and on the suction surface of the blade are also considerably reduced.
Unsteady Rotor Heat Transfer in a Transonic Turbine Stage
Authors: Didier F. , De´nos R. , Arts T.
Published: February 2009

This experimental investigation reports the convective heat transfer coefficient around the rotor of a transonic turbine stage. Both time-resolved and time-averaged aspects are addressed. The measurements are performed around the rotor blade at 15%, 50% and 85% span as well as on the rotor tip and the hub platform. Four operating conditions are tested covering two Reynolds numbers and three pressure ratios. The tests are performed in the compression tube turbine test rig CT3 of the von Karman Institute, allowing a correct simulation of the operating conditions encountered in modern aero-engines. The time-averaged Nusselt number distribution shows the strong dependence on both blade Mach number distribution and Reynolds number. The time-resolved heat transfer rate is mostly dictated by the vane trailing edge shock impingement on the rotor boundary layer. The shock passage corresponds to a sudden heat transfer increase. The effects are more pronounced in the leading edge region. The increase of the stage pressure ratio causes a stronger vane trailing edge shock and thus larger heat transfer fluctuations. The influence of the Reynolds number is hardly visible.
Effect of Secondary Flow on Heat Transfer Over a Guide Vane Suction Side
Authors: Khalatov Artem , Syred Nick , Khalatov Sergei , Bowen Philip
Published: February 2009

An accurate design of a guide vane has become a major issue in gas turbine engineering since secondary flow in vane cascade has a significant effect on the end wall and airfoil suction surfaces. Heat transfer and average flow measurements were performed in a linear guide vane cascade consisting of five gas turbine vanes. All measurements were taken on the suction surface of the middle (third) airfoil. The height-to-pitch ratio H/S was varied from 1.0 to 0.17, the inlet boundary layer thickness δin was either 2 mm or 10 mm. The vortex structures influence the average flow parameters, the laminar-to-turbulence transition, and the local heat transfer. This is especially pronounced in the area close to the end wall area. The inlet flow temperature non-uniformity has no influence on the heat transfer, however local flow temperature ‘spikes’ were registered at x = 0.6 due to the flow enthalpy migration in the span-wise direction. A few useful correlations are suggested as a result of the experimental data processing.
St and cf Augmentation for Real Turbine Roughness With Elevated Freestream Turbulence
Authors: Bons Jeffrey P.
Published: February 2009

Experimental measurements of skin friction (cf) and heat transfer (St) augmentation are reported for low speed flow over turbine roughness models. The models were scaled from surface measurements taken on actual, in-service land-based turbine hardware. Model scaling factors ranged from 25 to 63, preserving the roughness height to boundary layer momentum thickness ratio for each case. The roughness models include samples of deposits, TBC spallation, erosion, and pitting. Measurements were made in a zero pressure gradient turbulent boundary layer at two Reynolds numbers (Rex = 500,000 and 900,000) and three freestream turbulence levels (Tu = 1%, 5%, and 11%). Measurements at low freestream turbulence indicate augmentation factors ranging from 1.1–1.5 for St/Sto and from 1.3–3.0 for cf /cfo (Sto and cfo are smooth plate values). For the range of roughness studied (average roughness height, k, less than 1/3rd the boundary layer thickness) the level of cf augmentation agrees well with accepted equivalent sandgrain (ks) correlations when ks is determined from a roughness shape/density parameter. This finding is not repeated with heat transfer, in which case the ks-based St correlations overpredict the measurements. Both cf and St correlations severely underpredict the effect of roughness for k+ < 70 (when ks, as determined by the roughness shape/density parameter, is small). A new ks correlation based on the rms surface slope angle overcomes this limitation. Comparison of data from real roughness and simulated (ordered cones or hemispheres) roughness suggests that simulated roughness is fundamentally different from real roughness. Specifically, ks values that correlate cf for both simulated and real roughness are found to correlate St for simulated roughness but overpredict St for real roughness. These findings expose limitations in the traditional equivalent sandgrain roughness model and the common use of ordered arrays of roughness elements to simulate real roughness surfaces. The elevated freestream turbulence levels produce augmentation ratios of 1.24 & 1.5 (St/Sto) and 1.07 & 1.16 (cf /cfo) compared to the Tu = 1% flow over the smooth reference plate. The combined effects of roughness and elevated freestream turbulence are greater than their added effects suggesting that some synergy occurs between the two mechanisms. Specifically, skin friction augmentation for combined turbulence and roughness is up to 20% greater than that estimated by adding their separate effects and 8% greater than compounding (multiplying) their separate effects. For heat transfer augmentation, the combined effect of turbulence and roughness is 5% higher than that estimated by compounding their separate effects at high freestream turbulence (Tu = 11%). At low turbulence (Tu = 5%), there is a negative synergy between the two augmentation mechanisms as the combined effect is now 13% lower than that estimated by compounding their separate effects.
The Measurement of Full-Surface Internal Heat Transfer Coefficients for Turbine Airfoils Using a Non-Destructive Thermal Inertia Technique
Authors: Nirmalan Nirm V. , Bunker Ronald S. , Hedlung Carl R.
Published: February 2009

A new method has been developed and demonstrated for the non-destructive, quantitative assessment of internal heat transfer coefficient distributions of cooled metallic turbine airfoils. The technique employs the acquisition of full-surface external surface temperature data in response to a thermal transient induced by internal heating/cooling, in conjunction with knowledge of the part wall thickness and geometry, material properties, and internal fluid temperatures. An imaging Infrared camera system is used to record the complete time history of the external surface temperature response during a transient initiated by the introduction of a convecting fluid through the cooling circuit of the part. The transient data obtained is combined with the cooling fluid network model to provide the boundary conditions for a finite element model representing the complete part geometry. A simple 1D lumped thermal capacitance model for each local wall position is used to provide a first estimate of the internal surface heat transfer coefficient distribution. A 3D inverse transient conduction model of the part is then executed with updated internal heat transfer coefficients until convergence is reached with the experimentally measured external wall temperatures as a function of time. This new technique makes possible the accurate quantification of full-surface internal heat transfer coefficient distributions for prototype and production metallic airfoils in a totally non-destructive and non-intrusive manner. The technique is equally applicable to other material types and other cooled/heated components.
Computational Analysis of Pitch-Width Effects on the Secondary Flows of Turbine Blades
Authors: Xu C. , Amano R. S. , Marini B.
Published: February 2009

A three-dimensional computational code was developed for solving time-averaged flows within a turbine blade row using a novel time-marching method. A concept of incorporating dissipation terms into the time derivative terms was proposed to allow the code to have the capability of handling both incompressible and compressible flows. The code was validated by comparing the computational results with experiments in a turbine stator blade passage. The code was further used to investigate the influence of secondary flow in a turbine blade row due to different pitch-width ratios. Detailed secondary flows as well as loss profiles in different sizes of root pitch-width ratio are presented and discussed. The results of this study provide useful information for evaluation of the secondary flow effects due to the pitch-width ratio influence for the future new turbine blade designs.
An Experimental Investigation on the Influence of Periodic Unsteady Inflow Conditions on Leading Edge Film Cooling
Authors: Wolff Stefan , Fottner Leonhard , Ardey Sabine
Published: February 2009

In order to close the gap in knowledge concerning the influence of periodic unsteady inflow conditions on the mixing process of film cooling jets, time resolved flow velocities and turbulent fluctuations in the injection zone of a linear large scale high pressure turbine cascade with leading edge film cooling were measured by means of the 3D hot-wire anemometry. The periodic impinging wakes are generated by a wake generator consisting of moving bars upstream of the cascade inlet plane. Near the stagnation point the blades are equipped with one row of film cooling holes on the suction side and one row on the pressure side. Mach number and Reynolds number are set to values typically found in modern gas turbines. At a density ratio of unity the blowing ratio is set to M = 0.7 and the Strouhal number is set to Sr = 0.31. The general flow structures which were determined by the steady state measurements — i.e. normal jet in cross flow behavior with the kidney shape vortex on the suction side and a second counter rotating pair of vortices underneath the kidney vortex on the pressure side — have been detected by the unsteady measurements as well. The large recirculation zone behind the pressure side injection hole, caused by the strong adverse pressure gradient and the lift off of the coolant jet, is not suppressed by the passing wakes but rather reinforced promoting potential hot spots in this area. The suction side coolant jet has almost disappeared when the wake leaves the suction side region. It undergoes a recovery process until the next wake hits again on the suction side.
Characterization of the Velocity and Heat Transfer Fields in an Internal Cooling Channel With High Blockage Ratio
Authors: Casarsa Luca , C¸akan Murat , Arts Tony
Published: February 2009

The present experimental study is dealing with a detailed aero/thermal investigation of the turbulent flow inside a rib-roughened turbine blade cooling channel by means of Particle Image Velocimetry (PIV) and Liquid Crystal Thermometry (LCT). The main objectives of the paper are to provide detailed information about the behaviour of such a complicated flow, useful for its understanding, and to create a wide and reliable data base for numerical code validation. The measurements are carried out at engine representative Reynolds number within a scaled up model of a stationary straight cooling channel, with turbulent promoters (or ribs) installed on one wall. The ribs have an angle of attack of 90 deg with respect to the “mean” flow direction; their blockage ratio is equal to 30%. Detailed wall heat transfer distributions are presented. The main time-averaged flow features are identified and quantified; a number of rms characteristics are put in evidence and compared to the heat transfer distributions.
Turbulent Heat Transfer Measurements on a Wall With Concave and Cylindrical Dimples in a Square Channel
Authors: Moon S. W. , Lau S. C.
Published: February 2009

Dimpled surfaces may be considered for heat transfer enhancement in internal cooling of gas turbine airfoils. In this study, convective heat transfer and pressure drop for turbulent airflow in a square channel with a dimpled wall were examined. Experiments were conducted to determine the average heat transfer coefficient on the dimpled wall and the overall pressure drop across the channel, for nine concave and cylindrical dimples with various diameters and depths, and for Reynolds numbers (based on the channel hydraulic diameter) between 10,000 and 65,000. For the concave and cylindrical dimple configurations studied, the dimples were found to enhance the heat transfer coefficient by 70% (1.7 times) to over three times the value for fully developed turbulent flow through a smooth tube, with increase of the overall pressure drop of over four times. For both the concave and cylindrical dimples, heat transfer was enhanced more when the dimples covered a larger portion of the surface of the wall. The cylindrical dimples caused higher overall heat transfer coefficient (based on the projected area) and lower pressure drop than the concave dimples with the same diameters and depths. Thus, cylindrical dimple configuration may be a better alternative than concave dimples in enhancing heat transfer, for the experimental conditions and dimple configurations investigated. Further experiments are recommended to determine if cylindrical dimples of other dimensions also give higher thermal performances than concave dimples of the same dimensions, subjected to other flow and thermal boundary conditions, such as irregular channels with or without rotation.
Computation of Turbulent Heat Transfer on the Walls of a 180 Degree Turn Channel With a Low Reynolds Number Reynolds Stress Model
Authors: Rigby D. L. , Ameri A. A. , Steinthorsson E.
Published: February 2009

The Low Reynolds number version of the Stress-ω model and the two equation k-ω model of Wilcox were used for the calculation of turbulent heat transfer in a 180 degree turn simulating an internal coolant passage. The Stress-ω model was chosen for its robustness. The turbulent thermal fluxes were calculated by modifying and using the Generalized Gradient Diffusion Hypothesis. The results showed that using this Reynolds Stress model allowed better prediction of heat transfer compared to the k-ω two equation model. This improvement however required a finer grid and commensurately more CPU time.
Heat Transfer in a Complex Trailing Edge Passage for a High Pressure Turbine Blade: Part 1 — Experimental Measurements
Authors: Bunker Ronald S. , Wetzel Todd G. , Rigby David L.
Published: February 2009

A combined experimental and computational study has been performed to investigate the detailed heat transfer coefficient distributions within a complex blade trailing edge passage. The experimental measurements are made using a steady liquid crystal thermography technique applied to one major side of the passage. The geometry of the trailing edge passage is that of a two-pass serpentine circuit with a sharp 180-degree turning region at the tip. The upflow channel is split by interrupted ribs into two major sub-channels, one of which is turbulated. This channel has an average aspect ratio of roughly 14:1. The spanwise extent of the channel geometry includes both area convergence from root to tip, as well as taper towards the trailing edge apex. The average section Reynolds numbers tested in this upflow channel range from 55,000 to 98,000. The tip section contains a turning vane near the extreme corner. The downflow channel has an aspect ratio of about 5:1, and also includes convergence and taper. Turbulators of varying sizes are included in this channel also. Both detailed heat transfer and pressure distribution measurements are presented. The pressure measurements are incorporated into a flow network model illustrating the major loss contributors.
Spatially-Resolved Heat Transfer and Flow Structure in a Rectangular Channel With 45° Angled Rib Turbulators
Authors: Mahmood G. I. , Ligrani P. M. , Won S. Y.
Published: February 2009

Spatially-resolved Nusselt numbers and flow structure are presented for a stationary channel with an aspect ratio of 4 and angled rib turbulators inclined at 45° with perpendicular orientations on two opposite surfaces. The flow structure results include time-averaged distributions of streamwise velocity and total pressure, surveyed over flow cross-sectional planes, as well as flow visualization images and friction factors. Results are given at different Reynolds numbers based on channel height from 270 to 90,000. The ratio of rib height to hydraulic diameter is .078, the rib pitch-to-height ratio is 10, and the blockage provided by the ribs is 25 percent of the channel cross-sectional area. Spatially-resolved local Nusselt numbers are highest on tops of the rib turbulators, with lower magnitudes on flat surfaces between the ribs, where regions of flow separation and shear layer re-attachment have pronounced influences on local surface heat transfer behavior. Also important are intense, highly unsteady secondary flows and vortex pairs, which increase secondary advection and turbulent transport over the entire channel cross-section. The resulting augmented local and spatially-averaged Nusselt number ratios (rib turbulator Nusselt numbers normalized by values measured in a smooth channel) generally increase on the rib tops as Reynolds number increases. Nusselt number ratios decrease on the flat regions away from the ribs, especially at locations just downstream of the ribs, as Reynolds number increases. Globally-averaged Nusselt number ratios vary from 3.36 to 2.82 as Reynolds number increases from 10,000 to 90,000. Thermal performance parameters also decrease somewhat as Reynolds number increases over this range, with values in approximate agreement with, or slightly higher than 60° continuous rib data measured by other investigators in a square channel.
A Numerical Study of Flow and Heat Transfer in Rotating Rectangular Channels (AR=4) With 45° Rib Turbulators by Reynolds Stress Turbulence Model
Authors: Al-Qahtani Mohammad , Chen Hamn-Ching , Han Je-Chin
Published: February 2009

Computations were performed to study three-dimensional turbulent flow and heat transfer in a rotating smooth and 45° ribbed rectangular channels for which heat transfer data were available. The channel aspect ratio (AR) is 4:1, the rib height-to-hydraulic diameter ratio (e/Dh) is 0.078 and the rib-pitch-to-height ratio (P/e) is 10. The rotation number and inlet coolant-to-wall density ratios, Δρ/ρ, were varied from 0.0 to 0.28 and from 0.122 to 0.40, respectively, while the Reynolds number was fixed at 10,000. Also, two channel orientations (β − 90° and 135° from the rotation direction) were investigated with focus on the high rotation and high density ratios effects on the heat transfer characteristics of the 135° orientation. These results show that, for high rotation and high density ratio, the rotation induced secondary flow overpowered the rib induced secondary flow and thus change significantly the heat transfer characteristics compared to the low rotation low density ratio case. A multi-block Reynolds-Averaged Navier-Stokes (RANS) method was employed in conjunction with a near-wall second-moment turbulence closure. In the present method, the convective transport equations for momentum, energy, and turbulence quantities are solved in curvilinear, body-fitted coordinates using the finite-analytic method.
Effect of Rotation on Flow and Temperature Distributions in a Two-Pass Channel Connected by Two Rows of Holes
Authors: Pramanick Achintya , Ekkad Srinath V.
Published: February 2009

This paper presents a computational study of the effect of rotation on the velocity and thermal field for a two-pass channel connected by two rows of holes on the divider wall. Detailed velocity and temperature distributions are presented inside a rotating two-pass coolant channel connected by two rows of holes on the divider walls. The enhanced cooling in this passage design is achieved by a combination of impingement and crossflow-induced swirl. The cross flow is generated from one coolant passage to the adjoining coolant passage through a series of straight or angled holes placed along the dividing wall. The holes deliver the flow from one passage to another typically achieved in a conventional design by an 180° U-bend. In this study, the configuration has two rows of 12 straight holes placed axially. The two hole rows are placed at 1.27-cm distance from the sidewalls and are angled orthogonally. There are 12 holes in each rows of 1.27-cm diameter each. The holes direct flow perpendicular to the axial direction. Commercial software, FLUENT, is used for predicting the flow using the standard k-ε turbulence model. The results are presented at a channel flow Reynolds number of 25000. The effect of rotation number from 0, 0.1, and 0.2 is studied along with inlet coolant-to-wall density ratio of 0.05, 0.1, and 0.15. Results show that the impingement and swirl flow are affected by rotation induced Coriolis and centrifugal forces. There appears to be little effect of centrifugal buoyancy for this geometry as velocity profiles are unaffected by the wall temperature changes.