bluuebunny/crossref_metadata_2025 huggingface dataset:
Vaneless Diffuser Advanced Model (2005D)
Authors: Dubitsky Oleg , Japikse David
Published: November 2008

A new vaneless diffuser model is presented. Upon thorough examination of the change in average total pressure, the average static pressure and the average flow angle through a vaneless diffuser, it was discovered that existing models fail to provide useful integrity. Consequently, a new model was built. It was learned that it was necessary to use a two-zone model of the flow entering the vaneless diffuser and to carefully model the two-zone degradation as the flow passes through the vaneless diffuser. The new model is presented with detailed testing. The impact upon future design is outlined, and the expectation is established that various future designs will require the integrity of the new model; old models can be used in limited cases with care.
Cooling Performance of the Steam-Cooled Vane in a Steam Turbine Cascade
Authors: Bohn Dieter , Ren Jing , Kusterer Karsten
Published: November 2008

With the increasing demand of the electricity, an efficiency improvement and thereby a reduced CO2 emission of the coal-fired plants are expected in order to reach the goal of the Kyoto protocol. In the medium term, increase of thermal efficiency of the steam cycle can be achieved by a rise of the process parameters. Currently, live steam pressures and temperatures up to 300 bar and 923 K are planned as a next step. This means an increase of 50–100 K in comparison to plants presently under operation. Such a large increase will only be possible, if all thermally high loaded components of the plant are made of new materials and/or new technologies for cooling and thermal insulation are applied. Closed circuit steam cooling of blades and vanes in modern steam turbines is a promising technology in order to establish elevated live steam temperatures in future steam turbine cycles. Due to the cooling of the components, materials can be used, which are common in todays application. In this paper, a steam-cooled test vane in a cascade with external hot steam flow is analysed numerically by applying a Conjugate Calculation Technique. The boundary conditions of the hot steam are derived from the planned inlet parameter of the HP steam turbine in a medium term and a long term in order to show the application potential of such a cooling technique. The supply steam parameters of the cooling are set equal to those of the outlet steam from the HP steam turbine. The results show that the steam cooling technique can reach necessary cooling effectiveness in the steam turbine by using only minor amounts of cooling steam. The internal part of the vane is cooled homogeneously in both investigated cases. Basic features and phenomena of the steam cooling technique in steam turbines are discussed in the paper with respect to future application. It has been found that the trailing edge region of the steam-cooled vane is less critical with respect to high thermal load than in gas turbine application with the same configuration. At the leading edge of the vane, high thermal load can be detected because of the direct impingement of the hot steam. Improvement of the leading edge cooling is recommended by increasing cooling steam amount for the leading edge channel. Thus, a larger diameter of the channel can increase internal cooling efficiency for the leading edge. Due to the different heat transfer characteristics on the suction side and pressure side, the suction side shows also a higher temperature distribution in the investigated cases.
The Interaction of Turbine Inter-Platform Leakage Flow With the Mainstream Flow
Authors: Reid Kevin , Denton John , Pullan Graham , Curtis Eric , Longley John
Published: November 2008

Individual nozzle guide vanes (NGVs) in modern aero engines are often cast as a single piece with integral hub and casing endwalls. When in operation there is a leakage flow through the chord-wise inter-platform gaps. An investigation into the effect of this leakage flow on turbine performance is presented. Efficiency measurements and NGV exit area traverse data from a low speed research turbine are reported. Tests show that this leakage flow can have a significant impact on turbine performance, but that below a threshold leakage fraction this penalty does not rise with increasing leakage flow rate. The effect of various seal clearances are also investigated. Results from steady-state simulations using a three-dimensional multiblock RANS solver are presented with particular emphasis paid to the physics of the mainstream/leakage interaction and the loss generation.
The Use of a Circumferentially Nonuniform Stator to Attenuate Harmful Aerodynamic and Mechanical Interactions in an Advanced Mixed Flow Splittered Rotor/Tandem Variable Stator LP Compressor
Authors: Barton M. T. , Gentile D. P.
Published: November 2008

A potential flow computer model that can handle blade row interaction problems has been used to analyze the circumferential static pressure distribution at the trailing edge plane of an advanced mixed flow splittered rotor low pressure compressor produced by a downstream tandem stator/strut system. The computer model is based on the Douglas-Neumann formulation and features a powerful automated optimization feature which can define a restagger pattern that will either minimize stator blade-to-blade loading differences or minimize the circumferential static pressure nonuniformity on a preselected axial plane. The latter approach was used presently to design a circumferentially nonuniform stagger angle distribution for the second row of the tandem stator that reduced the circumferential static pressure variation, and consequently its harmonic excitation, on the upstream rotor. Comparison between the predicted static pressure distribution for the baseline case (with uniform stagger angles) and the optimized design showed a dramatic reduction in the circumferential pressure variation. Fourier analyses of the pressure distributions confirmed a substantial decrease in the magnitude of the harmonic index thought to be responsible for the observed severe rotor vibration. The analytical results were confirmed by back-to-back engine tests of the baseline and restaggered tandem stators, where light probe measurements of splitter blade synchronous response showed a proportional decrease in rotor response for the optimized stator configuration. The restagger was implemented quickly and easily using simple modifications to the stator unison ring. With the high splitter blade deflections reduced to an acceptable level, the engine was able to safely accelerate through the former prohibited speed range and continue with the planned test sequence. The restaggered stator design and test effort, from initial problem statement through a successful demonstration test in the gas generator, was accomplished in less than 6 weeks.
Unsteady Flow Investigations and Their Recent Challenges in Compressor Design
Authors: Peitsch Dieter , Poensgen Christian , Mu¨ck Bernhard
Published: November 2008

This paper describes the relevance of unsteady flow investigations in turbomachinery and how they are included into the design process for compressors of modern aero engines. Significant work has been performed in leadership of Professor Gallus at the Institute for Jet Propulsion and Turbomachinery (IST) at the RWTH Aachen, University of Technology, Germany, in this area of interest. Selected topics of this numerical and experimental work will be shown. This academic work has of course to be made useful for application in the design process of turbomachinery, so the connection to today’s challenges is shown as well. In recent years, blisks have become a more attractive choice for design and manufacture of compressor rotors. In contrast to the earlier years, where blade flutter has enjoyed the main focus of interest, the blisk technology requires different approaches in terms of the assessment of the interaction between structure and fluid. The implementation of the flutter and forced response investigations into the design process for blisks is thus a major issue to be solved. The paper will describe, how the relevant unsteady flow phenomena are assessed. Results will be shown for a recent engine development, the high pressure compressor for the new TP400 engine for the Airbus A400M military transport aircraft. This compressor is currently designed at Rolls-Royce Deutschland (RRD).
A Contribution to the Abrasive Effect of Particles in a Gas Turbine Pre-Swirl Cooling Air System
Authors: Schneider O. , Benra F.-K. , Dohmen H. J. , Jarzombek K.
Published: November 2008

With the increase of cooling air passing through the internal air system of modern gas turbines, a greater number of airborne particles is transported to the film cooling holes in the turbine blade surface. In spite of their small size, these holes are critical for airflow and must be free of blockage. A test rig has been designed to study the quantity of separated particles at various critical areas of the internal air system. Former publications for this conference gave detailed insight into the test rig, the flow structure and the particle motion during separation. The process of separation generates abrasion on the rotating and stationary parts of the system. When considering service and maintenance or even unexpected operation faults of the gas turbine, it is important to know the location and abrasion rate of these critical areas. The flow structure within the pre-swirl cooling air system results in locally focused abrasion regions, which are investigated in this paper. New simulations, taking additional physical effects into account, are discussed in the paper. The simulation results are compared to results obtained by measurements and observations within the test rig. Qualitative and quantitative results show the ability to predict the quantity of abrasion during operation on various critical areas of the system.
PIV Maps of Tip Leakage and Secondary Flow Fields on a Low Speed Turbine Blade Cascade With Moving Endwall
Authors: Palafox P. , Oldfield M. L. G. , LaGraff J. E. , Jones T. V.
Published: November 2008

New, detailed flow field measurements are presented for a very large low-speed cascade representative of a high-pressure turbine rotor blade with turning of 110 degrees and blade chord of 1.0 m. Data was obtained for tip leakage and passage secondary flow at a Reynolds number of 4.0 × 105, based on exit velocity and blade axial chord. Tip clearance levels ranged from 0% to 1.68% of blade span (0% to 3% of blade chord). Particle Image Velocimetry (PIV) was used to obtain flow field maps of several planes parallel to the tip surface within the tip gap, and adjacent passage flow. Vector maps were also obtained for planes normal to the tip surface in the direction of the tip leakage flow. Secondary flow was measured at planes normal to the blade exit angle at locations upstream and downstream of the trailing edge. The interaction between the tip leakage vortex and passage vortex is clearly defined, revealing the dominant effect of the tip leakage flow on the tip endwall secondary flow. The relative motion between the casing and the blade tip was simulated using a motor-driven moving belt system. A reduction in the magnitude of the under-tip flow near the endwall due to the moving wall is observed and the effect on the tip leakage vortex examined.
Analysis of Flow Mechanism in 2-D Compressor Cascade With Splitter
Authors: Li Haipeng , Liu Houxing , Zou Zhengping
Published: November 2008

Through numeric simulation of flow in different turning angle cascades with different splitters, the flow mechanism in 2-D compressor cascade with splitter is particularly analyzed in this paper. The study result shows that the splitter can influence the flow field in cascade intensively; the positive attack angle on the splitter leading is an important cause that brings about cascade loss growth. If the splitter chord is short, the presence of splitter has little influence on the cascade inlet flow field, but has intense influence on the outlet flow field. by the modification of splitter, the pressure distribution of cascade can be greatly changed, the separation flow in the cascade can be restrained effectively, and the performance of splitter cascade can be effectively improved. On the condition of maintaining a high mass flux, the presence of splitter can greatly raise the loading of cascade.
Unsteady Numerical Simulation of Shock Systems in Vaneless Counter-Rotating Turbine
Authors: Wang Huishe , Zhao Qingjun , Zhao Xiaolu , Xu Jianzhong
Published: November 2008

A detailed unsteady numerical simulation has been carried out to investigate the shock systems in the high pressure (HP) turbine rotor and unsteady shock-wake interaction between coupled blade rows in a 1+1/2 counter-rotating turbine (VCRT). For the VCRT HP rotor, due to the convergent-divergent nozzle design, along almost all the span, fishtail shock systems appear after the trailing edge, where the pitch averaged relative Mach number is exceeding the value of 1.4 and up to 1.5 approximately (except the both endwalls). A group of pressure waves create from the suction surface after about 60% axial chord in the VCRT HP rotor, and those waves interact with the inner-extending shock (IES). IES first impinges on the next HP rotor suction surface and its echo wave is strong enough and cannot be neglected, then the echo wave interacts with the HP rotor wake. Strongly influenced by the HP rotor wake and LP rotor, the HP rotor outer-extending shock (OES) varies periodically when moving from one LP rotor leading edge to the next. In VCRT, the relative Mach numbers in front of IES and OES are not equal, and in front of IES, the maximum relative Mach number is more than 2.0, but in front of OES, the maximum relative Mach number is less than 1.9. Moreover, behind IES and OES, the flow is supersonic. Though the shocks are intensified in VCRT, the loss resulted in by the shocks is acceptable, and the HP rotor using convergent-divergent nozzle design can obtain major benefits.
Flow Field and Heat Transfer Characteristics in an Impingement/Effusion Cooling System for Combustor Liner
Authors: Xu Quanhong , Zhang Chi , Lin Yuzhen , Liu Gaoen
Published: November 2008

The present study is conducted to investigate the characteristics of the flow field and heat transfer in an impingement/effusion cooling scheme for gas turbine combustor liner. It is designed to provide an insight, through the study of the flow field, into the physical mechanisms responsible for the enhanced impingement heat transfer near the effusion hole entrance. In this impingement/effusion cooling scheme, the angle between the impingement hole and effusion hole and the wall surface are 90 deg and 30 deg respectively. The square arrays of impingement/effusion holes are used with equal numbers of holes offset half a pitch relative to each plate so that an impingement jet is located on the center of each four effusion holes and vice versa. The flow field of the double skin wall space is described by the way of Particle Image Velocimetry (PIV). Two kinds of target plates, with and without effusion holes, are used in the impingement heat transfer study. Through changing the impingement Reynolds and the impingement gap, the change of the impingement heat transfer coefficient on the target plates is investigated. The impingement heat transfer test results show that the impingement heat transfer is enhanced near the entrance of the effusion holes, which could fully explain the feature of the impingement heat transfer coefficient on the target plate.
An Experimental and Numerical Study of Labyrinth Seal Flow
Authors: Vakili A. D. , Meganathan A. J. , Michaud M. , Radhakrishnan S.
Published: November 2008

Leakage flow in 2-D constant rotor diameter stepped labyrinth seals is investigated by means of pressure and velocity field measurements and numerical simulation of 2-D and axisymmetric models. The basis of investigation is a generic stepped labyrinth seal currently used in industry in steam turbine generators. The performance of the baseline seal design was compared with new seal designs with specific features changed in order to examine their influence on leakage characteristics through such seals. Numerical modeling and experiments were performed over a range of seal pressure ratios from 1 to 10. A number of configurations were evaluated both experimentally and numerically. This paper discusses flow details associated with only one configuration as compared with the baseline. Results have been helpful in the understanding of seal flow leakage and total pressure loss mechanisms. Mechanisms of leakage reduction in labyrinth seals included turbulence induced viscous losses, chamber vortex generation, flow stagnation losses, and increased flow streamline curvature. Numerical results provided insight into the flow field details and were helpful in facilitating basic physical understandings used for improved seal designs.
Influence of Laser Drilling Imperfection on Film Cooling Performances
Authors: Jovanovic´ Milenko B. , de Lange Hendrik C. , van Steenhoven Anton A.
Published: November 2008

Film cooling is studied on a jet in a cross-flow. The influence of a hole production imperfection on the jet-cross flow interaction is investigated experimentally by means of particle image velocimetry and liquid crystal thermography. To simulate the imperfection a half torus was placed inside the hole. The experiments were conducted with and without the production imperfection and the velocity ratio was varied. If the imperfection is absent, vortices are generated by means of the Kelvin-Helmholtz instability and separation on the hole trailing edge. The imperfection produces additional vortical structures and the flow field starts to oscillate. The deteriorated flow field changes the heat transfer. Surface temperature measurements show that the production imperfection reduces the film cooling effectiveness. The influence of the production imperfection on the film cooling effectiveness decreases with the enhanced velocity ratio.
CFD Analysis of a 15 Stage Axial Compressor: Part I — Methods
Authors: Belamri T. , Galpin P. , Braune A. , Cornelius C.
Published: November 2008

The flow field of a 15 stage axial compressor is analyzed using a 3-D Navier-Stokes CFD tool. The compressor under investigation is a prototype first compressor version before optimization of the Siemens V84.3A family. A methodology is described for steady state and transient flow simulations of the entire 15 stages compressor in one computation (not piece by piece). The simulation includes tip gaps, mass bleeds, hub leakage flows, and ranges from single passage to full 360 degrees analysis. The work is divided into two companion papers. This first part, “CFD Analysis of a 15 Stage Axial Compressor Part I: Methods” describes the overall methodology used, based on a middle portion of the compressor R5S9 (Rotor 5 to Stator 9). Various effects are analyzed: mesh style and refinement, boundary conditions, steady or transient, tip clearance, and numerical issues (turbulence model choice, advection model choice, parallel processing performance). A high sensitivity of the predictions to the tip clearance height was found. Excellent design point predictions are obtained with steady-state frame change interface models (Stage average interface), as well as with transient simulations (transient rotor-stator interface).
Shape Optimization of Turbine Stage Using Adaptive Range Differential Evolution and Three-Dimensional Navier-Stokes Solver
Authors: Song Liming , Feng Zhenping , Li Jun
Published: November 2008

A new optimization method named as Adaptive Range Differential Evolution (ARDE) is proposed and developed for the turbine stage design. The mathematical tests are used to demonstrate the optimization performance of the present ARDE through compared with the Simple Genetic Algorithms (SGA) and the Differential Evolution (DE). Combined with the ARDE, surface modeling method and Navier-Stokes solver, a low aspect ratio transonic turbine stage is optimized, with 28 design variables in total, for the maximization of the isentropic efficiency. The optimization design of this case is performed on the cluster parallel Personal Computers. The optimal design turbine stage shows a better aerodynamic performance than that of the reference design while meeting the strength requirement. The robustness and reliability of the presented ARDE for the turbomachinery optimization design are also illustrated.
Effects of Corrugation Angle on Flow and Heat Transfer in Cross Corrugation Channels With Sinusoidal Waves
Authors: Yin Jixiang , Li Guojun , Feng Zhenping
Published: November 2008

Fluid field and heat transfer in a cross-corrugated (CC) geometry with sinusoidal wavy plates have been investigated with laminar flow assumption. The 7 × 7 multiple channels configuration with six different corrugation intersection angles θ are studied. For air mass flow rates in the range of 50 ≤ Re ≤ 2000 and P/Hi = 2.2, average Nu and Fanning friction factor f are analyzed. Numerical results show that Fanning friction factors increase with corrugation intersection angle θ. But the Re of transition to turbulence decreases when θ > 90°. Moreover, the absence of a sharp transition is demonstrated for θ ≤ 90°. The distribution of local Nu becomes slightly flatter along the width when θ increases. The maximal Nu locates near the main flow trailing position and at the entrances of the neighborhood outflows, the minimum Nu locates at the contact points of top plate and bottom plate. Average Nu shows different variety with various ranges of corrugation intersection angle θ and Re, the maximum average Nu is obtained with 120° plate. Predictive correlations of f and Nu with Re have been carried out base on the numerical results. The sinusoidal wavy plates have a Nu which is about 6.5 times higher than that of a parallel-plate channel. And the concomitant f is about 54 times higher than that of the sinusoidal wavy plates. At last, the comprehensive assessment of the surfaces performance shows that the 30° plate surface has the optimal performance, while the one with 150° angle has the worst performance.
Numerical Modeling of Flow and Thermal Patterns Within a Combustor Simulator
Authors: Kunze V. R. , Wolff M. , Barringer M. D. , Thole K. A. , Polanka M. D.
Published: November 2008

Performance enhancements and control of heat transfer in high pressure gas turbine vanes and rotors is dependent on understanding the flow and thermal fields approaching the turbine. The flow field exiting the combustor has highly non-uniform pressure and temperature variations in both the radial and circumferential directions as well as high turbulence levels. Several studies have shown significant impact on the overall and secondary flow fields within the turbine due to the inlet profile. The Turbine Research Facility (TRF) at Wright-Patterson Air Force Base has recently added a non-reactive full scale annular combustor simulator to the facility to study these effects. In conjunction with the TRF experimental effort to simulate combustor section exit flows, a three-dimensional CFD analysis of the newly installed simulator has been undertaken. The analysis aids in the experimental implementation of the simulator and gives further understanding of the simulator’s complex internal flow patterns. The goal for the TRF simulator is to produce a wide range of profiles to the inlet plane of the vane for evaluation of effects on heat transfer, loss, and loading. The CFD analysis allows an understanding of how those profiles are obtained by tracking the flow through two rows of staggered dilution holes and six rows of staggered film cooling holes on both the ID and OD liners of the main simulator chamber. This enables control as the CFD can guide the experimenter in knowing which liner component influenced the turbine inlet profile shape. Cases can then be run computationally by varying the mass flows and temperatures to tailor the profile to the desired shape prior to running the experiment. These profiles can then be sent through the turbine stage both computationally and experimentally to understand their impact. Finally, turbine airfoils and cooling patterns can then be designed to take advantage of this knowledge.
Dynamic Modeling and Simulation of Distributed Parameter Heat Exchanger
Authors: Zhang Huisheng , Weng Shilie , Su Ming
Published: November 2008

The heat exchanger is one of the key components for an advanced gas turbine power plant. This paper presents a new modeling technique to solve the iterative calculation for the coupling of pressure and mass flow rate in the flow system of a distributed parameter heat exchanger. In this modeling method, a control volume is separated into a volume unit and a flow resistance unit. The volume unit takes into account the flow rate difference between the inlet and outlet of the control volume, and the flow resistance unit considers the pressure drop in the control volume. Thus a single set of ordinary differential equations is obtained. When we adopt this modeling technique, a non-iterative dynamic model of heat exchanger can be established based on conservation theorems. A single plate-fin heat exchanger for the inter-cooled recuperated gas turbine engine is taken as an example to study the steady state and dynamic performance of the heat exchanger and to verify this modeling technique. The simulation results are reasonable, and the results satisfy the requirement of the real time model. This provides a powerful tool for the dynamic simulation and optimization design of advanced gas turbine power system.
Pre-Swirled Cooling Air Delivery System Performance
Authors: Chew J. W. , Ciampoli F. , Hills N. J. , Scanlon T.
Published: November 2008

This paper reports results from numerical simulations of the flow in pre-swirl cooling air delivery systems. Two different systems have been investigated corresponding to experimental rigs for which measured data is available. The rigs are representative of aero-engine conditions. The difference in the performance of the two rigs has been addressed. The flow in the pre-swirl nozzles and in the pre-swirl chambers has been investigated separately. For the pre-swirl chamber a simplified model, in which the nozzle is replaced by a slot, has been used to reduce the computational effort required. Nevertheless numerical results are in good agreement with experimental measurements. It is shown that the difference in the geometry of the pre-swirl chambers is largely responsible for the difference in performance of the rigs. Numerical results have also been compared with predictions from a previously published simplified model. An adjustment of the empirical constants in the simple model has been proposed in order to improve the prediction of the moments in the pre-swirl chamber.
Numerical Prediction of Film Cooling and Heat Transfer on the Leading Edge of a Rotating Blade With Two Rows Holes in a 1-1/2 Turbine Stage at Design and Off Design Conditions
Authors: Yang Huitao , Chen Hamn-Ching , Han Je-Chin , Moon Hee-Koo
Published: November 2008

Numerical simulations were performed to predict the film cooling effectiveness and the associated heat transfer coefficient on the leading edge of a rotating blade in a 1-1/2 turbine stage using a Reynolds stress turbulence model together with a non-equilibrium wall function. Simulations were performed for both the design and off-design conditions to investigate the effects of blade rotation on the leading edge film cooling effectiveness and heat transfer coefficient distributions. It was found that the tilt stagnation line on the leading edge of rotor moves from the pressure side to the suction side, and the instantaneous coolant streamlines shift from the suction side to the pressure side with increasing rotating speed. This trend was supported by the experimental results. The result also showed that the heat transfer coefficient increases, but film cooling effectiveness decreases with increasing rotating speed. In addition, the unsteady characteristics of the film cooling and heat transfer at different time phases, as well as different rotating speeds, were also reported.
Flow Measurements in a First Stage Nozzle Cascade Having Endwall Contouring, Leakage and Assembly Features
Authors: Piggush J. D. , Simon T. W.
Published: November 2008

This work supports new gas turbine designs for improved performance by evaluating the use of endwall contouring in a cascade that is representative of a first stage stator passage. Contouring accelerates the flow, reducing the thickness of the endwall inlet boundary layer to the turbine stage and reducing the strength of secondary flows within the passage. Reduction in secondary flows leads to less mixing in the endwall region. This allows improved cooling of the endwall and airfoil surfaces with injected and leakage flows. The present paper documents component misalignment and leakage flow effects on the aerodynamic losses within a passage having one contoured and one straight endwall. Steps on the endwall and leakage flows through the endwall can lead to thicker endwall boundary layers, stronger secondary flows and possibly additional vortex structures in the passage. The paper compares losses with steps of various geometries and leakage of various flow rates to assess their importance on aerodynamic losses in this contoured passage. In particular, features associated with the combustor-to-turbine transition piece and the slashface gap, a gap between two vane segments on the vane platform, are addressed. An n-factorial study is used to quantify the importance of such effects on aerodynamic losses.
CFD for Jet Impingement Heat Transfer With Single Jets and Arrays
Authors: Ibrahim Mounir B. , Kochuparambil Bejoy J. , Ekkad Srinath V. , Simon Terrence W.
Published: November 2008

CFD experiments were conducted for heat transfer with jet impingement over solid surfaces. The parameters include: 1) Jet Reynolds number from 3,000 to 23,000, 2) Jet-to-target-plate spacing (z/d), from 2 to 14 (single jet), d is jet diameter, 3) Target plate shape: 3a) flat, 3b) concave, 3c) convex, (single jet), 4) One row of seven jets impinging on a flat surface, the channel has one end closed (at 24d away from the most upstream jet axis), 5) Three rows of seven jets each in-line arrangement impinging on a flat surface, the channel has one end closed (at 24d away from the most upstream jet axis). Four CFD models (utilizing FLUENT commercial code) have been considered: 1) laminar flow (no turbulent transport), and turbulent flow with turbulence modeling by 2) the standard k–ε model, 3) the k–ω model, and 4) the v2–f model. The predictions of Nu number for each case were compared with experimental data available from the literature. It is shown that the v2–f model gives the best overall performance, though the k–ω model gives good predictions for most of the flow, with the exception of near the stagnation zone for some cases. The models are in much better agreement (with the data) as z/d grows and at larger radial locations from the jet axis, as expected. For multiple jets in one row (z/d = 2), again the v2–f showed the best overall agreement with the experimental data. The k–ω model is not as good while k–ε clearly overpredicts the Nusselt numbers. For multiple jets in three inline rows (z/d = 5), all the three models were in overall agreement with the experimental data. However, k–ε and k–ω exhibit an important phenomenon, reported by the experiments: a decrease of the stagnation Nu from the upstream jet to the downstream ones. The v2–f model did not reproduce this feature.
Film Cooling Effectiveness on the Leading Edge of a Rotating Film-Cooled Blade Using Pressure Sensitive Paint
Authors: Ahn Jaeyong , Schobeiri M. T. , Han Je-Chin , Moon Hee-Koo
Published: November 2008

Detailed film cooling effectiveness distributions were measured on the leading edge region of a rotating blade using a Pressure Sensitive Paint technique. The film cooling effectiveness information was obtained from the oxygen concentration difference between air and nitrogen or air and CO2 injection cases by applying the mass transfer analogy. The blowing ratio was controlled to be 0.5, 1.0, and 2.0 while the density ratios of 1.0 and 1.5 were obtained using nitrogen and CO2 as coolant gases, respectively. Tests were conducted on the first stage rotor of a 3-stage axial turbine at 2400, 2550, and 3000 rpm. The Reynolds number based on the axial chord length and the exit velocity was 200,000 and the total to exit pressure ratio was 1.12 for the first rotor. The film cooling effectiveness distributions were presented along with the discussions on the influences of blowing ratio, density ratio, and vortices around the leading edge region at different rotational speeds.
Spatial Arrangement Dependance of Cooling Performance of an Integrated Impingement and Pin Fin Cooling Configuration
Authors: Nakamata Chiyuki , Okita Yoji , Matsuno Shinsuke , Mimura Fujio , Matsushita Masahiro , Yamane Takashi , Fukuyama Yoshitaka , Yoshida Toyoaki
Published: November 2008

Experimental and numerical studies were conducted for the development of the integrated impingement and pin-fin cooling configuration. In the development, the spatial arrangements of impingement hole, pin-fin and film cooling (discharge) hole were the main concern. The temperature measurement was performed for different test pieces with various spatial arrangements to clarify the cooling effectiveness variation with the arrangement and the other cooling parameters. Experiments were conducted with 673K hot gas flow and room temperature cooling air. The Reynolds number of gas side flow was 380000 and cooling air Reynolds number was 5000–30000. Test plate surface temperatures were measured using an infrared camera. The cooling effectiveness obtained from the experiment for one specimen was different from that for a specimen that had the same pin density but a different spatial arrangement. So it was confirmed that an arrangement of hole and pin, as well as pin density, was an important parameter. CFD analysis was also conducted to make clear how spatial arrangement affected internal heat transfer characteristics. Pressure losses were also evaluated for each specimen, and total thermal performance was compared. A basic configuration with one pin at the center of a unit area showed the most superior total thermal performance.
Impingement Cooling for Modern Combustors: Experimental Analysis and Preliminary Design
Authors: Facchini Bruno , Surace Marco , Tarchi Lorenzo
Published: November 2008

The aim to reach very low emission limits has recently changed several aspects of combustor fluid dynamics. Among them, combustor cooling experienced significant design efforts to obtain good performances with unfavorable conditions. This paper deals with experimental research and 1D numerical simulations of impingement cooling from multiple jet arrays, performed in the first two years of the European research project LOPOCOTEP. Geometries are derived from typical LPP combustor cooling configurations, i.e. small and sparse holes, due to low coolant mass flow rate and high pressure losses (compared to typical blade cooling parameters). Tests are performed with both transient and steady-state technique, using TLC (Thermochromic Liquid Crystal) wide-band formulations. Results are obtained in terms of local distributions of heat transfer coefficient and effectiveness. Average values are compared with calculations to validate numerical code and derive useful indications on effectiveness variation for different blowing rate. It is accordingly found that literature results almost match the measurements. It is also shown that measured row-by-row effectiveness values can be usefully employed in a preliminary design stage. Data concerning holes discharge coefficients are also presented.
Numerical Heat Transfer Analysis of an Innovative Gas Turbine Combustor: Coupled Study of Radiation and Cooling in the Upper Part of the Liner
Authors: Andreini A. , Bacci A. , Carcasci C. , Facchini B. , Asti A. , Ceccherini G. , Del Puglia E. , Modi R.
Published: November 2008

A numerical study of a single can combustor for the GE10 heavy-duty gas turbine, which is being developed at GE-Energy (Oil & Gas), is performed using the STAR-CD CFD package. The topic of the present study is the analysis of the cooling system of the combustor liner’s upper part, named “cap”. The study was developed in three steps, using two different computational models. As first model, the flow field and the temperature distribution inside the chamber were determined by meshing the inner part of the liner. As second model, the impingement cooling system of the cold side of the cap was meshed to evaluate heat transfer distribution. For the reactive calculations, a closure of the BML (Bray-Moss-Libby) approach based on Kolmogorov-Petrovskii-Piskunov theorem was used. The model was implemented in the STAR-CD code using its user coding features. Then the radiative thermal load on the liner walls was evaluated by means of the STAR-CD-native Discrete Transfer model. The selection of the radiative properties of the flame was performed using a correlation procedure involving the total emissivity of the gas, the mean beam length and the gas temperature. The estimated heat flux on the cap was finally used as boundary condition for the calculation of the cooling system, consisting of 68 staggered impingement jet lines on the cold side of the cap. The resulting temperature distribution shows a good agreement with the experimental values measured by thermocouples. The results confirm the validity of the implemented procedure, and point out the importance of a full CFD computation as an additional tool to support classic correlation design procedures.
Turbine Blade Platform Impingement Cooling With Effects of Film Extraction and Roughness
Authors: Bunker Ronald S. , Bailey Jeremy C.
Published: November 2008

With the movement of combustion systems to flatter exit temperature profiles for the purpose of reduced emissions, vane endwalls and blade platforms are experiencing higher gas temperatures than previously. While vane endwalls may be cooled by conventional methods extended from the airfoil techniques, blade platforms require simple but robust cooling methods that are amenable to the high rotational loads. This study presents a platform cooling solution that utilizes a single impingement jet directed to the underside of the platform from the forward shank face. Detailed local distributions of heat transfer coefficients are presented for the cooled platform region formed by the cavity between the pressure side and suction side of two adjacent blade castings. Fundamental understanding is provided for impingement into this confined cavity region, with two differing jet positions. The effects of jet impingement distance to the surface from Z/D of 1.5 to 8, and jet Reynolds number from 65,000 to 155,000 are investigated. Furthermore, to augment the cooled surface heat transfer coefficients, surface roughness is applied by a patented method appropriate to actual turbine hardware. The single cooling jet produces a platform cooling distribution that may be treated as two distinct regions. The primary impingement region is seen to behave in line with expectations from prior literature, with some cavity geometry effects. Impingement heat transfer coefficients are increased by as much as 50% through the use of an applied randomly roughened surface that models an actual braze alloy method. This roughness benefit is however lost if the impingement cooling jet is not relatively close to the platform surface. The non-impingement heat transfer regions of the platform are affected only by the total cooling flow rate supplied by the impingement jet.
The Influence of Turbulence on Wake Dispersion and Blade Row Interaction in an Axial Compressor
Authors: Henderson Alan D. , Walker Gregory J. , Hughes Jeremy D.
Published: November 2008

The influence of free-stream turbulence on wake dispersion and boundary layer transition processes has been studied in a 1.5-stage axial compressor. An inlet grid was used to produce turbulence characteristics typical of an embedded stage in a multistage machine. The grid turbulence strongly enhanced the dispersion of inlet guide vane (IGV) wakes. This modified the interaction of IGV and rotor wakes, leading to a significant decrease in periodic unsteadiness experienced by the downstream stator. These observations have important implications for the prediction of clocking effects in multistage machines. Boundary layer transition characteristics on the outlet stator were studied with a surface hot-film array. Observations with grid turbulence were compared with those for the natural low turbulence inflow to the machine. The transition behavior under low turbulence inflow conditions with the stator blade element immersed in the dispersed IGV wakes closely resembled the behavior with elevated grid turbulence. It is concluded that with appropriate alignment, the blade element behavior in a 1.5-stage axial machine can reliably indicate the blade element behavior of an embedded row in a multistage machine.
Novel Unsteady Temperature/Heat Transfer Instrumentation and Measurements in the Presence of Combustor Instabilities
Authors: Chana K. S. , Syed K. J. , Wedlock M. I. , Copplestone R. W. , Cook M. S. , Bulat G.
Published: November 2008

Lower NOx emissions from gas turbine combustion systems can be achieved through reducing the equivalence ratio of the main reaction zone and/or increasing the burner pressure drop. This strategy however takes pressure drop and/or air away from the combustor cooling, thereby compromising the combustor life. In order to achieve an optimum design that is a good compromise between low emissions and long component life, accurate heat transfer prediction is essential. It is well known that free stream turbulence can influence wall heat transfer characteristics. However the impact of combustion induced pressure dynamics, and the associated unsteady fluid dynamics, upon combustor wall heat transfer has not been adequately investigated. This paper reports on combustion tests conducted at gas turbine operating conditions, where pressure dynamics have been controlled by altering combustor operating conditions and through the use of a siren placed in the upstream air flow. Combustor wall temperatures were measured using standard thermocouples and QinetiQ’s “True Surface Thermocouples” (TST). The latter, which were mounted on the hot gas surface of the wall, are capable of a fast response and are capable of indicating the temperature fluctuations experienced by the metal surface. Fourier analysis of the TSTs showed no particular peaks associated with the pressure dynamics. This suggests that any coherence is damped within the boundary layer or by the thermal inertia of the metal. However temperature fluctuations of up to about 100°C were detected.
Rotating Stall Characteristics in a Wide Vaneless Diffuser
Authors: Ljevar S. , de Lange H. C. , van Steenhoven A. A.
Published: November 2008

Paper reports a numerical study on vaneless diffuser flow instability performed for the purpose of better understanding of rotating stall mechanism in radial vaneless diffusers. This analysis is restricted to the two-dimensional flow where effect of wall boundary layers is neglected. Numerical results reveal that a two-dimensional rotating flow instability similar to rotating stall occurs when critical flow angle is exceeded. They also show that the stability limit and the structure of a two dimensional rotating instability are influenced by the configuration geometry and inlet and outlet flow conditions. Good agreement with data from the literature is found for the stability limit and number and speed of propagating cells. Number of cells and their speed is somewhat higher than observed in experiments from literature. This might imply that inception point is caused by the core flow instability and that wall boundary layers are more determinative for the structure of rotating instability.
The Importance of Shroud Leakage Modelling in Multistage Turbine Flow Calculations
Authors: Rosic Budimir , Denton John D. , Pullan Graham
Published: November 2008

Three-dimensional steady multistage calculations, using the mixing plane approach, are compared with experimental measurement in a low speed 3-stage model turbine. The comparisons are made with two levels of shroud seal clearance, one representative of a real turbine and one with minimal seal clearance and almost no shroud leakage. Three different calculations are compared. The first computes the main blade path with no modelling of shroud leakage. The second includes a simple model of shroud leakage using sources and sinks on the end-walls and the third is a multiblock calculation with all leakage paths and cavities computed. It is found that neglect of shroud leakage makes the computed velocity profiles and loss distributions significantly different to those measured. Simple modelling of shroud leakage gives some improvement but full calculation of the leakage flows and cavities is necessary to obtain good agreement between calculation and measurement.
A Pressure Gradient Sensitive Wall Function for the Prediction of Turbulent Flow in Thermal Turbomachinery
Authors: Pieringer Paul , Sanz Wolfgang
Published: November 2008

The economical design of thermal turbine stages of high efficiency requires very fast CFD tools of high accuracy. Since efficiency is mainly determined by secondary flow, a major focus in CFD is laid on boundary layer treatment. By resolving the boundary layer using a fine grid, usually the best results can be achieved, but at cost of computational time and memory. Wherever resolving the boundary layer is unacceptable, wall functions are used for simulating the flow close to solid walls. This paper describes a novel wall function approach for the prediction of three-dimensional turbulent boundary layer flows. The model is designed especially for the application to thermal turbomachinery in order to properly predict accelerated, decelerated and separated boundary layers at transonic conditions. The model is based on the law-of-the-wall expression by Spalding and also considers the shear-stress distribution perpendicular to the wall surface, which is mainly determined by pressure gradients. To cope with the boundary conditions, this wall function model uses phantom (ghost) cells, that must be located within the boundary layer, so there is a clearance between the first calculation cell and the wall. This allows an improved grid resolution for wall function approaches. To verify the accuracy of the model, results are shown for turbomachinery vane test cases at both steady state and time-resolved flow conditions.
Review and Understanding on Sweep in Axial Compressor Design
Authors: Ji Lucheng , Chen Jiang , Lin Feng
Published: November 2008

As a means of improving axial compressor performance, sweep technique has been investigated for over half a century and gained wide uses, in the past one decade. However, there is still diverse controversy about the roles of sweep in axial compressor design. In this paper, historical remarks about the sweep are presented firstly. Then, an understanding about the role of sweep is put forward. That is, the sweep is a degree of freedom (DOF) of blade design that emphasizes on matching the aerodynamic loading of every blade element along the whole span within the full operation range. The present understanding about the role of sweep may lead it a more sophisticated use.