bluuebunny/crossref_metadata_2025 huggingface dataset:
Local Heat/Mass Transfer Characteristics on a Rotating Blade With Flat Tip in a Low Speed Annular Cascade: Part 2 — Tip and Shroud
Authors: Rhee Dong-Ho , Cho Hyung-Hee
Published: November 2008

The local heat/mass transfer characteristics on the tip and shroud were investigated using a low speed rotating turbine annular cascade. Time-averaged mass transfer coefficients on the tip and shroud were measured using a naphthalene sublimation technique. A low speed wind tunnel with a single stage turbine annular cascade was used. The turbine stage is composed of sixteen guide plates and blades. The chord length of blade is 150 mm and the mean tip clearance is about 2.5% of the blade chord. The tested Reynolds number based on inlet flow velocity and blade chord is 1.5×105 and the rotational speed of the blade is 255.8 rpm at design condition. The results were compared with the results for a stationary blade and the effects of incidence angle of incoming flow were examined for incidence angles ranging from −15 to +7 degree. The off-design test conditions are obtained by changing the rotational speed with a fixed incoming flow velocity. Flow reattachment on the tip near the pressure side edge dominates the heat transfer on the tip surface. Consequently, the heat/mass transfer coefficients on the blade tip are about 1.7 times as high as those on the blade surface and the shroud. However, the heat transfer on the tip is about 10% lower than that for the stationary case due to reduced leakage flow with the relative motion. The peak regions due to the flow reattachment are reduced and shifted toward the trailing edge and additional peaks are formed near the leading edge region with decreasing incidence angles. But, quite uniform and high values are observed on the tip with positive incidence angles. The time-averaged heat/mass transfer on the shroud surface has similar a level to that of the stationary cases.
Radial Migration of Shed Vortices in a Transonic Rotor Following a Wake Generator: A Comparison Between Time Accurate and Average Passage Approaches
Authors: Turner Mark G. , Gorrell Steven E. , Car David
Published: November 2008

This paper shows a comparison of an unsteady simulation using Turbo and an Average Passage simulation for a two blade row configuration consisting of a wake generator followed by a transonic rotor. Two spacings were simulated, both close and far. The unsteady results compare well with experiment especially for the profile of efficiency difference between close and far. An analysis of results helps to explain the unusual profile seen experimentally that is due to the radial migration of wake generator shed vortices with negative radial velocities near the tip. In addition, different components of the average passage body forces (deterministic stresses) are explored that shows the main terms are the axial momentum and the metal blockage.
Modelling of Film Cooling: Part I — Experimental Study of Flow Structure
Authors: Bernsdorf Stefan , Rose Martin G. , Abhari Reza S.
Published: November 2008

This paper, which is Part I of a two part paper, reports on new experimental data taken in a steady flow, flat plate wind tunnel at ETH Zu¨rich, while Part II utilizes this data for calibration and validation purpose of a new film cooling model embedded in a 3D CFD code. 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 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 three dimensional flow structures are revealed.
The Design, Development and Evaluation of 3D Aerofoils for High Speed Axial Compressors: Part 1 — Test Facility, Instrumentation and Probe Traverse Mechanism
Authors: Lippett D. , Woollatt G. , Ivey P. C. , Timmis P. , Charnley B. A.
Published: November 2008

This paper, in two parts, reports measurements from, and simulation of, Cranfield University’s 3-stage high-speed axial compressor. Using this newly built rig, supported by European Commission, a consortium of gas-turbine companies have tested a set of conventionally stacked 2D rotor and stator blades. The results from this experiment were used to evaluate and assess the performance of several commercially available CFD codes leading to the collaborative design of an advanced three-dimensional blade set seeking, if possible, a 2% efficiency gain. The limited axial spacing between the measurement planes and the blade rows required the design of a unique seven probe assembly and traverse mechanism able to yaw and pitch the probes and to control the insertion depths. This mechanism was designed to accommodate different probes, such as cobra, fast response (pneumatic) and temperature measuring probes, and deliver area traverses between rotor and stators throughout the compressor. For probe calibration a high speed wind tunnel section was designed to accommodate this mechanism enabling calibrations for Mach numbers up to 0.78, as well as for a wide range of pitch and yaw angles values. This mechanism combined with a post processing programme incorporating a mapping technique for the relative offset of the measurement points on the probe secured very detailed results throughout the compressor. Measurements show the complex three dimensional flow structure and secondary flows associated with tip-leakage, endwall boundary layers, wake transportation and blade row interactions. The importance of a rigorous mapping procedure was particularly useful where the wake thickness was small and pressure gradients high in comparison to the probe size.
Stator Clocking Effects on 3D Flow in a Two-Stage Low-Pressure Turbine
Authors: Krysinski Jan E. , Blaszczak Jaroslaw R. , Smolny Antoni
Published: November 2008

Detailed experimental investigations were conducted of the detailed flow structures in a model two-stage low-pressure turbine. The objective of this study is to gain more insight into phenomena affecting flow behaviour due to the indexing airfoil effects in multi-stage machines. To investigate the effect of the vane clocking on axial turbine efficiency and performance, an analysis of the experimental results of the steady and unsteady flow measurements throughout the flow field for different circumferential positions of the first stator is performed. Detailed flow pictures were measured for all measuring planes according to the radial and circumferential inter-row traversing. Aerodynamic measurements were compared to the turbine design speed and “classical” clocking position of the two stators. Herein, some steady and unsteady flow measurement results, and the external characteristics for different circumferential positions of the stator vanes are described. The experimental investigations presented in this paper were carried out since mid-nineties [1] on a 2-stage turbine research facility TM-3 at the Institute of Turbomachinery of Technical University of Lodz in close cooperation with Institute of Jet Propulsion and Turbomachinery of Technical University of Aachen, Germany.
Prediction of Heat Transfer in a Ribbed Channel: Evaluation of Unsteady RANS Methodology
Authors: York William D. , Holloway D. Scott , Leylek James H.
Published: November 2008

Heat transfer in a straight channel with rib turbulators on one wall is predicted numerically with an unsteady Reynolds-averaged Navier-Stokes (URANS) methodology and compared to code-validation quality experimental data from the literature. Additionally, for comparison, steady simulations of the problem are conducted using two popular turbulence closure models, a Realizable k-ε model and a differential Reynolds-stress model. Closure in the URANS simulation is provided by a new eddy-viscosity-based model that was developed in the Advanced Computational Research Laboratory at Clemson University. This new model consists of three transport equations, and it is designed specifically to promote natural unsteadiness in the flow without the need for artificial forcing. In all cases, the Reynolds number, based on hydraulic diameter, is equal to 24,000. Eight square ribs, orthogonal to the flow direction, are equally spaced on the bottom wall of the channel. For the URANS simulation, after the flow becomes fully-developed in the streamwise direction, the predicted Nusselt number on the ribbed wall follows the trend of measured data from the modeled experimental study. However, the unsteady simulation slightly overpredicts the distance to the peak heat transfer aft of each rib. Also, the heat transfer prediction is very dependent on the grid resolution aft of the ribs. Therefore, efficient refinement of the unstructured mesh and grid-independence issues are discussed. Results of both steady simulations show a significant underprediction of Nusselt number over the entire ribbed wall, with the Reynolds-stress model giving the better result of the two steady closure models. The results of this study clearly show that unsteady vortex shedding off of the ribs is important in the physics of this problem, and a systematic, unsteady methodology is necessary to accurately predict ribbed-channel heat transfer.
Analytical and Computational Study of Radial Loads in Volutes and Collectors
Authors: Chochua Gocha , Koch Jay M. , Sorokes James M.
Published: November 2008

The physical mechanism of the radial load generation in discharge volutes and collectors of a centrifugal compressor is analytically studied. A new equation is presented enabling prediction of radial load magnitude and direction based on compressor geometry and operational conditions. The analytical approach is compared with computational fluid dynamics (CFD) results utilizing different rotor-stator interface models and test data measurements. The effect of a low solidity vaned diffuser on radial load attenuation is presented.
Heat Transfer and Pressure Investigation of Dimple Impingement
Authors: Kanokjaruvijit K. , Martinez-Botas R. F.
Published: November 2008

Heat transfer and pressure results of an inline array of round jets impinging on a staggered array of dimples are reported with the consideration of various geometric and parametric effects; results are normalized against flat plate data. The heat transfer was measured by using transient wideband liquid crystal method. The geometrical configurations considered were crossflow (or spent-air exit) scheme, dimple geometries and impinging positions. Three crossflow schemes were tested such as one-way, two-way and free exits. These led to the idea of the coupling effects of impingement and channel flow depending on which one dominated. Hemispherical and cusped elliptical dimple shapes with the same wetted area were considered, and found that both dimples showed the similarity in heat transfer results. Impinging positions on dimples and on flat portions adjacent to dimples were examined. Throughout the study, the pitch of the nozzle holes was kept constant at 4 jet diameters. The investigated parameters were Reynolds number (ReDj) ranged from 5000 to 11 500, jet-to-plate spacing (H/Dj) varied from 1 to 12 jet diameters, dimple depths (d/Dd) of 0.15, 0.25 and 0.29, and dimple curvature (Dj/Dd) of 0.25, 0.50 and 1.15. The shallow dimples (d/Dd = 0.15) improved heat transfer significantly by 70% at H/Dj = 2 compared to that of the flat surface, while this value was 30% for the deep ones (d/Dd = 0.25). The improvement also occurred to the moderate and high Dj/Dd. The total pressure was a function of ReDj and H/Dj when H/Dj < 2, but it was independent of the target plate geometry. The levels of the total pressure loss of the dimpled plates were not different from those of the flat surface under the same set-up conditions. Wall static pressure was measured by using static taps located across each plate. ReDj and H/Dj affected the level of the static pressure while the dimple depth influenced the stagnation peaks, and the crossflow scheme affected the shape of the peaks.
Computational Analysis of Tonal Noise Generated High-Bypass Ratio Fan Stage
Authors: Sipatov Aleksey M. , Usanin Michail V. , Avgustinovich Valery G. , Chuhlantseva Natalia O.
Published: November 2008

The paper shows the procedure of estimating different fan stage geometries from the point of the fan stage aerodynamic and acoustic efficiency by using CFX v.5.6 gas dynamics software. Two different fan stages were examined. The acoustic analysis was made based on unsteady pressure distribution along exit guide vanes. The unsteady pressure distribution was determined from 3-D calculations of rotor-stator interactions. An improved approach is suggested to estimate acoustic sources with the leading edge geometry considered. Three various grid models were analyzed to estimate the grid discretization influence on computational results. The hybrid approach was used to evaluate the acoustic contribution of fan stage rotor–stator interaction to a turbo-jet engine total noise level on first harmonic in a far field. This approach consists of three steps. The first step includes the solving of Navier-Stokes equations in fan stage, the second step includes the solving of linearized Euler’s equations (LEE) in a near field and, then, the third covers the calculation of Ffowcs Williams–Hawkings (FWH) integral in a far field. The first step marked the effect of strong attenuation of acoustics modes in the fan passage. The results of calculations in the far field were compared with experiment data.
A Transient Calorimeter Technique for Determining Regional Heat Transfer Coefficients in the Three-Temperature Flowfield at a Turbine Airfoil Leading Edge
Authors: Nowlin Scott R. , Gillespie David R. H. , Ireland Peter T. , Knoche Ralf , Kingston T. Robert
Published: November 2008

In this paper, the authors develop a novel method of obtaining regionally-averaged heat transfer coefficients in flowfields characterized by three temperatures using the well-known transient calorimeter technique. The technique is used to determine heat transfer in aluminum models of idealized turbine blade leading edges cooled through internal surface impingement, film cooling feed passages, and external convective film cooling. The external surface is subject to a stagnating mainstream crossflow. Importantly, the contributions to heating from the external flow and cooling from the internal flow can be separately resolved solely by heating the internal flow. Results for a basic showerhead geometry and an advanced intersecting-passage cooling configuration are presented for a range of internal and external Reynolds numbers. The intersecting-passage model shows little improvement in heat transfer coefficient over the showerhead for the flow conditions tested; however, the total cooling carried out is improved by the increase in exposed surface area. The technique’s uncertainties are fully assessed.
Large-Eddy Simulation of Unsteady Surface Pressure Over a LP Turbine Blade Due to Interactions of Passing Wakes and Inflexional Boundary Layer
Authors: Sarkar S. , Voke Peter R.
Published: November 2008

The unsteady pressure over the suction surface of a modern low-pressure (LP) turbine blade subjected to periodically passing wakes from a moving bar wake generator is described. The results presented are a part of detailed Large-Eddy Simulation (LES) following earlier experiments over the T106 profile for a Reynolds number of 1.6×105 (based on the chord and exit velocity) and the cascade pitch to chord ratio of 0.8. The present LES uses coupled simulations of cylinder for wake, providing four-dimensional inflow conditions for successor simulations of wake interactions with the blade. The three-dimensional, time-dependent, incompressible Navier-Stokes equations in fully covariant form are solved with 2.4×106 grid points for the cascade and 3.05×106 grid points for the cylinder using a symmetry-preserving finite difference scheme of second-order spatial and temporal accuracy. A separation bubble on the suction surface of the blade was found to form under the steady state condition. Pressure fluctuations of large amplitude appear on the suction surface as the wake passes over the separation region. Enhanced receptivity of perturbations associated with the inflexional velocity profile is the cause of instability and coherent vortices appear over the rear half of the suction surface by the rollup of shear layer via Kelvin-Helmholtz (K-H) mechanism. Once these vortices are formed, the steady-flow separation changes remarkably. These coherent structures embedded in the boundary amplify before breakdown while traveling downstream with a convective speed of about 37 percent of the local free-stream speed. The vortices play an important role in the generation of turbulence and thus to decide the transitional length, which becomes time-dependent. The source of the pressure fluctuations on the rear part of the suction surface is also identified as the formation of these coherent structures. When compared with experiments, it reveals that LES is worth pursuing as an understanding of the eddy motions and interactions is of vital importance for the problem.
Experimental Evaluation of Turbocharger Turbine Performance Under Pulsating Flow Conditions
Authors: Szymko S. , Martinez-Botas R. F. , Pullen K. R.
Published: November 2008

The steady and pulsating performance results of a turbocharger mixed-flow turbine are presented. The results are taken at an equivalent speed of 70% (42,000rpm) for a pulse frequency range of 20 to 80 Hz. All instantaneous parameters required for unsteady performance evaluation are measured and discussed. Significant improvements to the measurement of instantaneous actual power have been carried out. Large variations in the operating point of the turbine occur in each pulse cycle, a velocity ratio range of 0.43 to 1.28 is seen for a 20 Hz pulse, this range reduces as the pulse frequency increases and unsteady effects become more prominent. During periods of turbine freewheeling, negative efficiencies can arise due to momentum transfer from the turbine to the working gas, although detrimental to the efficiency the energy content in these regimes are low. The use of a modified Strouhal number (MSt.) and a pressure modified Strouhal number (PMSt.) has proved useful in assessing when the onset of unsteadiness of the flow will become significant, a value of 0.1 has been used as an appropriate limit to steadiness. The results suggest that for a typical engine speed range the rotor may be considered quasi-steady whilst the turbine stage is predominately operating in an unsteady regime. Inference from the experimental data would suggest it is adequate to capture the performance of a turbine under pulsating flow using a ‘quasi-steady’ model when the MSt. < 0.1, and a ‘filling and emptying’ code when a PMSt. < 0.1 and above this value a ‘wave action’ model is more appropriate.
Separation and Transition Control on an Aft-Loaded Ultra-High-Lift LP Turbine Blade at Low Reynolds Numbers: High-Speed Validation
Authors: Vera Maria , Zhang Xue Feng , Hodson Howard , Harvey Neil
Published: November 2008

This paper presents the second part of an investigation of the combined effects of unsteadiness and surface roughness on an aft-loaded ultra high lift low pressure turbine (LPT) profile at low Reynolds numbers. The investigation has been performed using low-speed and high-speed cascade facilities. The low speed and the high speed profiles have been designed to have the same normalized isentropic Mach number distribution. The low speed results have been presented in Part 1 of this paper. The current paper examines the effect of different surface finishes on an aft-loaded ultra-high-lift LPT profile at Mach and Reynolds numbers representative of LPT engine conditions. The surface roughness values are presented along with the profile losses under steady and unsteady inflow conditions. The results show that the use of a rough surface finish might reduce the profile loss. In addition, the results show that the same quantitative values of losses are obtained at high and low speed flow conditions. The latter proves the validity of the low speed approach for ultra high lift profiles for the case of an exit Mach number of the order of 0.64. Hot wire measurements were carried out to explain the effect of the surface finish on the wake induced transition mechanism.
Simulation of Heat Transfer From Swirling Impinging Jets
Authors: Sunden Bengt , Larocque Jerome
Published: November 2008

This paper presents a numerical investigation of heat transfer from impinging swirling jets. Various turbulence models are used. Comparison with experimental data is provided. Generally the V2F model was found to perform better than other models but for small nozzle to plate distances it was hard to maintain good accuracy. The difference in performance, in terms of the Nusselt number distribution on the impingment wall, between swirling and non-swirling impinging jets was clearly identified. The magnitude of the swirling motion has great influence on the flow field and the formation of recirculating zones and accordingly the heat transfer process.
The Influence of Variable Rotor-Stator Interaction on Boundary-Layer Development in an Axial Compressor
Authors: Griebel Achim , Seume Joerg R.
Published: November 2008

Boundary layer and free stream measurements were taken at midspan for the first stator of a two-stage low-speed axial research compressor with IGV. The flow pattern on the blade suction surface is described for three different operating points and the transition onset is investigated. Rotor blades were manipulated to exhibit a larger upstream and downstream flow effect and were used to investigate the correlation between downstream wake propagation and upstream potential effect. In this study, these manipulated blades were used to unambiguously identify the potential and wake effects and their propagation in space and time. The wake causes induced transition on the downstream stator. Natural transition occurs between wakes. The amplified potential effect of the manipulated blade propagates upstream even beyond the stator.
Film Flow Around Bearing Chamber Support Structures
Authors: Eastwick Carol , Huebner Keno , Azzopardi Barry , Simmons Kathy , Young Colin , Morrison Richard
Published: November 2008

Cooling of aero-engine bearing chamber walls is achieved by the through-flow of thin oil films, which are typically only a few millimeters thick. Support structures and other features disrupt both air and oil motion within such chambers, leading to localized variations in the oil film thickness, causing variation in local heat transfer coefficients. An experimental study has been undertaken to establish the range of conditions under which dry-out occurs upstream and downstream of an obstruction. In addition film thickness measurements around the obstruction have been taken at a set air flow rate for four different liquid flow rates. A regime map of liquid vs. air Reynolds numbers has been created showing where stable and fluctuating type I and II dry-outs occur.
Parametrical Study of Hydrodynamic Seal Using a 2D Design Code and Comparing With a 3D CFD Model
Authors: Zheng Xiaoqing
Published: November 2008

Hydrodynamic non-contacting face seals, characterized with small leakage and long life, have been used in almost every sector of industrial pump applications. However, its application to aerospace has been proven challenging mainly owing to the high-speed, large size and vastly varied operating conditions in an engine environment. Design optimization is critical for a successful application in aero-engine. Seals designed with optimized hydrodynamic grooves based on a 2D code have been qualified for production. For certain seal applications, the optimized hydrodynamic grooves can be very different from what is usually used in aerospace applications. In order to verify such extreme design and confirm the optimization procedure, a three-dimensional CFD model was developed. The influences of hydrodynamic features were studied with the 3D CFD model and compared with the solution of the two-dimensional design code. The 3D computation generally supports the 2D result, and the seal design was further validated through rig tests.
Compressibility Effects on the Two Generations of Unsteady Flow Types in Axial Flow Compressors
Authors: Zhen Xinqian , Zhou Sheng , Hou Anping , Xiong Jinsong
Published: November 2008

There occurred unsteady separated flows inside axial flow compressors, which was however not taken into consideration in the present aerodynamic design system. This discrepancy indicates that the potential underlying unsteady separated flows is yet to be explored, hence the present research team proposes the concept of two generations of unsteady flow types, i.e. Unsteady Natural Flow Type (UNFT) and Unsteady Cooperative Flow Type (UCFT). Numerical simulations are carried out in the present paper to study the compressibility effect on the unsteady cooperative flow type in axial flow compressors. The studies show that aerodynamic performances are remarkably enhanced by means of transforming the flow type from UNFT into UCFT by imposing unsteady excitations. In the case of 2D subsonic cascade, performances are greatly improved in a wide range of Ma number (Ma < 0.8) and the maximum relative reduction of the loss coefficient reaches 40.2%. In the case of 2D trans-supersonic cascade, positive effects can’t be captured. However, in the case of a 3D trans-supersonic single rotor, the adiabatic efficiency is increased from 87.0% to 90.2%.
Thermal State Analysis of Industrial Gas Turbine Blades
Authors: Vassiliev V. , Kostege V. , Schnedler P. , Trishkin A.
Published: November 2008

The correct prediction GT hot components metal temperatures is decisive for lifetime assessments including predictions of fatigue, creep, oxidation, and corrosion. For the lifetime assessments of new components (at design phase) the thermal analysis is usually based on empirical correlations and simplified calculation models. To mitigate the uncertainties of the models and boundary conditions broad safety margins are included in the calculations. Therefore, the lifetime predictions at design phase can deviate significantly from the real achievable life. Based on the results of testing, prototype and field experience the predictions can be improved by corrections of models. This paper describes the procedure for thermal state predictions of industrial turbine blades and calibration procedure based on tests and field experience. The calibration procedure comprises: • Analysis of field experience, including temperature mapping. • Correction of thermal design models. • Recalculation of thermal state for different conditions, and re -evaluation of lifetime. Such approach allows to mitigate the limited accuracy of today’s design technology and/or allows to develop the improvements measures (e.g. for service interval extension). The paper describes the separate models implemented in this approach, and provides the example of industrial GT blading analysis. On basis of this analysis the measures for service interval extension have been developed.
Experiments With Three Dimensional Passive Flow Control Devices on Low-Pressure Turbine Airfoils
Authors: Bohl Douglas G. , Volino Ralph J.
Published: November 2008

The effectiveness of three dimensional passive devices for flow control on low pressure turbine airfoils was investigated experimentally. A row of small cylinders was placed at the pressure minimum on the suction side of a typical airfoil. Cases with Reynolds numbers ranging from 25,000 to 300,000 (based on suction surface length and exit velocity) were considered under low freestream turbulence conditions. Streamwise pressure profiles and velocity profiles near the trailing edge were documented. Without flow control a separation bubble was present, and at the lower Reynolds numbers the bubble did not close. Cylinders with two different heights and a wide range of spanwise spacings were considered. Reattachment moved upstream as the cylinder height was increased or the spacing was decreased. If the spanwise spacing was sufficiently small, the flow at the trailing edge was essentially uniform across the span. The cylinder size and spacing could be optimized to minimize losses at a given Reynolds number, but cylinders optimized for low Reynolds number conditions caused increased losses at high Reynolds numbers. The effectiveness of two-dimensional bars had been studied previously under the same flow conditions. The cylinders were not as effective for maintaining low losses over a range of Reynolds numbers as the bars.
Design and Analysis of a High Stage Loading Five-Stage LP Turbine Rig Employing Improved Transition Modeling
Authors: Gier Jochen , Hu¨bner Norbert
Published: November 2008

Modern aircraft engine designs have to realize high performance while managing important weight and cost issues. For high bypass ratio engines this results in high demands for the low pressure turbine component. The combination of a low speed parameter due to large fan diameters with reduced number of turbine stages due to weight and cost considerations leads to high stage loading requirements. Realizing high turbine efficiencies at low Reynolds Number operation at high altitudes for these high stage loadings is a key for state-of-the-art turbofan engines. For the aerodynamic design, high stage loading factors directly translate into increased airfoil flow turning. These increased turning levels have to be realized in combination with elevated airfoil load in order to keep low airfoil counts. This is particularly challenging in combination and requires minimization of negative effects on the airfoil losses and secondary losses. In this paper the development of a five-stage turbine with high stage loading is discussed with focus on the aerodynamic design. A specially designed cascade test has been part of the turbine design, which showed the need for improvement of the transition modeling within the CFD code used for design and analysis. The derived modeling enhancement is then employed for the analysis of the five-stage rig, which was tested at the altitude test facility at Stuttgart University. The modeling achieved significant improvements in the quality of the numerical results. The analysis includes sensitivities to Reynolds number and a detailed view of the suction side flow.
Numerically Aided Design of a High-Efficient Casing Treatment for a Transonic Compressor
Authors: Wilke I. , Kau H.-P. , Brignole G.
Published: November 2008

The transonic NASA Rotor 37 is well known for the occurrence of rotating stall at operation points close to stall. Several experimental and numerical investigations have shown that the instabilities observed mainly origin near the blade tip. Due to this property this rotor is theoretically suitable for a successful application of casing treatments. The first issue of this paper is to numerically investigate the flow mechanisms leading to compressor stall at design and off-design conditions. The numerical simulations carried out demonstrate that at design speed the behaviour of tip leakage flow in combination with the angle of attack at the blade tips is the key factor which limits the flow stability. Whereas at off-design, a steep angle of attack near the blade tip causes vast flow separations at the blade’s suction side. The second issue of the paper is to design an adapted casing treatment for the NASA Rotor 37. This issue is very demanding since the design total pressure ratio is with 2.1 extremely high and therefore, high-loss blockage zones can be easily caused by a misdesigned casing treatment. Based on the observations made in the first part guidelines are suggested for designing a casing treatment which efficiently enhances the compressor’s flow stability at design and off-design conditions. Casing treatments geometries are presented which address these suggested guidelines. Three-dimensional time accurate CFD-simulations are carried out to verify these designed casing treatments.
Scaling of Guide Vane Coolant Profiles and the Reduction of a Simulated Hot Streak
Authors: Jenkins Sean C. , Bogard David G.
Published: November 2008

The turbine section of a gas turbine engine is subjected to a non-uniform temperature distribution in the gas flow from the combustor. Regions of elevated temperatures, known as “hot streaks,” subject the turbine airfoil to high heat loads. In this study, the reduction of hot streaks by coolant from a film cooled nozzle guide vane was experimentally evaluated. Experiments were conducted with an approach mainstream turbulence level of 20% to simulate actual turbine conditions. The coolant distributions downstream of the vane were measured for varying blowing ratios and varying coolant density, and scaling methods were found for variations in both parameters. For this study, the hot streak peak was positioned to impact the vane at the stagnation line. Measurements of the hot streak strength with coolant blowing showed as much as a 55% decrease in peak temperature compared with no coolant.
Heat Transfer and Aerodynamics of Turbine Blade Tips in a Linear Cascade
Authors: Newton P. J. , Krishnababu S. K. , Lock G. D. , Hodson H. P. , Dawes W. N. , Hannis J. , Whitney C.
Published: November 2008

Local measurements of the heat transfer coefficient and pressure coefficient were conducted on the tip and near tip region of a generic turbine blade in a five-blade linear cascade. Two tip clearance gaps were used: 1.6% and 2.8% chord. Data was obtained at a Reynolds number of 2.3 × 105 based on exit velocity and chord. Three different tip geometries were investigated: a flat (plain) tip, a suction-side squealer, and a cavity squealer. The experiments reveal that the flow through the plain gap is dominated by flow separation at the pressure-side edge and that the highest levels of heat transfer are located where the flow reattaches on the tip surface. High heat transfer is also measured at locations where the tip-leakage vortex has impinged onto the suction surface of the aerofoil. The experiments are supported by flow visualisation computed using the CFX CFD code which has provided insight into the fluid dynamics within the gap. The suction-side and cavity squealers are shown to reduce the heat transfer in the gap but high levels of heat transfer are associated with locations of impingement, identified using the flow visualisation and aerodynamic data. Film cooling is introduced on the plain tip at locations near the pressure-side edge within the separated region and a net heat flux reduction analysis is used to quantify the performance of the successful cooling design.
Degradation of Film Cooling Performance on a Turbine Vane Suction Side Due to Surface Roughness
Authors: Rutledge James L. , Robertson David , Bogard David G.
Published: November 2008

After an extended period of operation, the surfaces of turbine airfoils become extremely rough due to deposition, spallation, and erosion. The rough airfoil surfaces will cause film cooling performance degradation due to effects on adiabatic effectiveness and heat transfer coefficients. In this study, the individual and combined effects of roughness upstream and downstream of a row of film cooling holes on the suction side of a turbine vane have been determined. Adiabatic effectiveness and heat transfer coefficients were measured for a range of mainstream turbulence levels and with and without showerhead blowing. Using these parameters, the ultimate film cooling performance was quantified in terms of net heat flux reduction. The dominant effect of roughness was a doubling of the heat transfer coefficients. Maximum adiabatic effectiveness levels were also decreased significantly. Relative to a film cooled smooth surface, a film cooled rough surface was found to increase the heat flux to the surface by 30% to 70%.
On the Momentum and Thermal Structures of Turbulent Spots in a Favourable Pressure Gradient
Authors: Chong T. P. , Zhong S.
Published: November 2008

This paper represents the results from an experimental investigation of the flow physics behind the difference in the transition zone length indicated by the momentum boundary layer and thermal boundary layer parameters observed on the suction surfaces of gas turbine blades. The experiments were carried out on turbulent spots created artificially in an otherwise laminar boundary layer developing over a heated flat plate in a zero pressure gradient and a favourable pressure gradient. A specially designed miniature triple wire probe was used to measure the streamwise velocity U, transverse velocity component V and temperature T simultaneously during the passage of the spots. In this paper, the general characteristics of the ensemble-averaged velocity and temperature perturbations, rms fluctuations and the second moment turbulent quantities are discussed and the influence of favourable pressure gradient on these parameters is examined. When a favourable pressure gradient is present, unlike in the velocity boundary layer where significant velocity fluctuations (or Reynolds shear stress) occur both on the plane of symmetry and the spanwise periphery, high temperature fluctuations (or turbulent heat fluxes) are confined in the plane of symmetry. The difference in the levels of velocity/temperature fluctuations at these two locations gives an indication of the effectiveness of momentum/heat transfer across the span of the spots. The results of this study show that the heat transfer within a spot is inhibited more than that of the momentum transfer at the presence of a favourable pressure gradient. This phenomenon is expected to slow down the spanwise growth of turbulent spots in the transitional thermal boundary layer, leading to a longer transitional zone length indicated by the heat transfer parameters as reported in the literature.
Supersonic and Transonic Compressors: Past, Status and Technology Trends
Authors: Broichhausen Klaus D. , Ziegler Kai U.
Published: November 2008

Transonic compressors with high leading edge Mach numbers are today state of the art in gas turbine design. This holds true for both aero engines and stationary gas turbines. In the early days of highly loaded compressors, the development started from the ideas about supersonic compressors with very high stage pressure ratios. This historical development and the basic ideas are described. The contributions of H. E. Gallus to this development are specially referred to. On that basis, today’s transonic compressors with reduced loading have been developed. The characteristic physics and design features of recent compressors are discussed with respect to aerodynamics, performance, structure mechanics and production technology. This is also done in view of the ideas of the pioneers in this field. Future technology trends of these compressor types as well as new compressor types are presented in the last part of the paper.
Improved Profile Loss and Deviation Correlations for Axial-Turbine Blade Rows
Authors: Zhu Junqiang , Sjolander Steen A.
Published: November 2008

Empirical correlations continue to play an important role in the early stages of turbine design and in the optimizing of the gas path. The recent development of highly-loaded low-pressure (LP) turbines has extended the design space beyond the range of geometric and aerodynamic parameters on which the existing correlations for profile losses and deviation are based. With the aid of a large database of measured profile losses, including in-house results and recent cases from the open literature, the profile loss predictions of Kacker and Okapuu have been reviewed. As a result, a revised correlation is proposed that appears to capture much better the loss behavior of axial-turbine airfoils of recent design, including very highly-loaded LP turbine airfoils. The opportunity was also taken to extend the range of applicability of the correlation for trailing-edge deviation originally proposed by Islam and Sjolander in 1999.
The Dynamics of the Horseshoe Vortex and Associated Endwall Heat Transfer: Part I — Temporal Behavior
Authors: Praisner T. J. , Smith C. R.
Published: November 2008

Instantaneous flow topology and the associated endwall heat transfer in the leading-edge endwall region of a symmetric airfoil are presented. An experimental technique was employed that allowed the simultaneous recording of instantaneous particle image velocimetry flow field and thermochromic liquid-crystal-based endwall heat transfer data. The endwall flow is dominated by a horseshoe vortex that forms from reorganized impinging boundary layer vorticity. A corner vortex is shown to be a steady feature of the corner region, while a secondary vortex develops sporadically immediately upstream of the horseshoe vortex. The region upstream of the horseshoe vortex is characterized by a bimodal switching of the near-wall reverse flow, which results in quasi-periodic eruptions of the secondary vortex. The bimodal switching of the reverse flow in the vicinity of the secondary vortex is linked to the temporal behavior of the down-wash fluid on the leading edge of the foil. Frequency analysis of the flow field and endwall heat transfer data, taken together, indicate that the eruptive behavior associated with the horseshoe vortex occurs at a frequency that is essentially the same as the measured turbulence bursting period of the impinging turbulent endwall boundary layer.
The Dynamics of the Horseshoe Vortex and Associated Endwall Heat Transfer: Part 2 — Time-Mean Results
Authors: Praisner T. J. , Smith C. R.
Published: November 2008

Time-mean endwall heat transfer and flow-field data in the endwall region are presented for a turbulent juncture flow formed with a symmetric bluff body. The experimental technique employed allowed the simultaneous recording of instantaneous particle image velocimetry flow field data, and thermochromic liquid-crystal-based endwall heat transfer data. The time-mean flow field on the symmetry plane is characterized by the presence of primary (horseshoe), secondary, tertiary, and corner vortices. On the symmetry plane the time-mean horseshoe vortex displays a bimodal vorticity distribution and a stable-focus streamline topology indicative of vortex stretching. Off the symmetry plane, the horseshoe vortex grows in scale, and ultimately experiences a bursting, or breakdown, upon experiencing an adverse pressure gradient. The time-mean endwall heat transfer is dominated by two bands of high heat transfer, which circumscribe the leading edge of the bluff body. The band of highest heat transfer occurs in the corner region of the juncture, reflecting a 350% increase over the impinging turbulent boundary layer. A secondary high heat-transfer band develops upstream of the primary band, reflecting a 250% heat transfer increase, and is characterized by high levels of fluctuating heat load. The mean upstream position of the horseshoe vortex is coincident with a region of relatively low heat transfer that separates the two bands of high heat transfer.
Investigation of the Gas Flow in the Rolling Element Bearing Assembly of a Centrifugal Compressor
Authors: Cui Michael M.
Published: November 2008

Combined with the geometric features, the pressure differential and bearing motion define the gas flow through the rolling-element-bearing assembly of a centrifugal compressor. The gas flow field then affects the oil distribution and heat transfer characteristics of the assembly accordingly. Investigations of the refrigerant gas flow through the rolling element bearing assembly of a centrifugal compressor are presented. A series of cases are studied for different operating conditions. The analyses include the geometric details of the assembly, such as the shaft, races, cages, balls, oil feeding system, and surrounding components. Refrigerant R123 is used as the working fluid. Both detailed three-dimensional flow field features and integrated parameters are calculated. The interactions between bearing motion and the surrounding structures are characterized. The flow patterns inside the bearings are defined. These results help us gain an insight into the basic physics that governs the bearing internal mass and heat transfer. The data and techniques developed can be used to design and optimize bearing and oil supply systems for the improvement of lubrication and cooling efficiency.
Simulation of Film Cooling Enhancement With Mist Injection
Authors: Li Xianchang , Wang Ting
Published: November 2008

Cooling of gas turbine hot section components such as combustor liners, combustor transition pieces, turbine vanes (nozzles) and blades (buckets) is a critical task for improving the life and reliability of hot-section components. Conventional cooling techniques using air-film cooling, impingement jet cooling, and turbulators have significantly contributed to cooling enhancements in the past. However, the increased net benefits that can be continuously harnessed by using these conventional cooling techniques seem to be incremental and are about to approach their limit. Therefore, new cooling techniques are essential for surpassing these current limits. This paper investigates the potential of film cooling enhancement by injecting mist into the coolant. The computational results show that a small amount of injection (2% of the coolant flow rate) can enhance the cooling effectiveness about 30% ∼ 50%. The cooling enhancement takes place more strongly in the downstream region, where the single-phase film cooling becomes less powerful. Three different holes are used in this study including a 2-D slot, a round hole, and a fan-shaped diffusion hole. A comprehensive study is performed on the effect of flue gas temperature, blowing angle, blowing ratio, mist injection rate, and droplet size on the cooling effectiveness with 2-D cases. Analysis on droplet history (trajectory and size) is undertaken to interpret the mechanism of droplet dynamics.
Axial Location of Casing Treatment in Multistage Axial Flow Compressors
Authors: Zhu Junqiang , Wu Yanhui , Chu Wuli
Published: November 2008

Koch in Ref. 1 presented a procedure for estimating the maximum pressure rise potential of axial flow compressor stages. When used in conjunction with a stage stacking off-design performance prediction method, it can indicate which stage is likely to be responsible for stalling a multistage compressor. Based on Koch’s approach, it is believed that the casing over the stage that stalls first (critical stage) should be treated in order to prevent or reduce the possibility of stalling. With the help of a two-stage compressor test rig, experimental investigations on the stage stalling and the ideal axial location of the treated casing in the multistage environment were performed using a recessed circumferential slot casing treatment. Experimental results indicate that it is possible to predict the stage that stalls first using Koch’s engineering method. In addition, it was verified that if the casing treatment was located over the critical stage, the stall margin of the whole multistage compressor could be improved significantly.