bluuebunny/crossref_metadata_2025 huggingface dataset:
Solid Oxide Fuel Cell Hybrid Systems Fed by Liquid Fuels for Distributed Power Generation
Authors: Santin Marco , Traverso Alberto , Massardo Aristide
Published: August 2009

Solid oxide fuel cell hybridization with micro gas turbines is an attractive option for distributed power generation up to a few MW, allowing to obtain high efficiency and low pollutant emission. In this publication, a comparative thermoeconomic analysis of SOFC hybrid systems with methanol and kerosene fuel processors is presented. Methanol can be produced from renewable sources. Also, hybrid systems fuelled by methanol can achieve high efficiencies due to effective heat recovery from the exhaust gases in the low temperature reformer. Kerosene is representative of conventional liquid fossil fuels, and it is also a typical fuel for aerospace applications. A 500 kW class hybrid system was chosen for this analysis and the performance was calculated based on macroscopic component models. The results were obtained with WTEMP software, developed by the Thermochemical Power Group of the University of Genoa. The choice of the fuel processing strategy and the influence of the main design parameters on the thermoeconomic characteristics of hybrid systems were investigated.
A NURBS-Based Optimization Tool for Axial Compressor Cascades at Design and Off-Design Conditions
Authors: Briasco Giuseppe , Bruna Dario , Cravero Carlo
Published: August 2009

Optimization techniques based on evolutionary strategies have become a general procedure in the industrial and academic worlds for the aero-mechanical design of turbomachinery blades. The airfoil geometry, parameterized using NURBS curve model, has been optimized minimizing the loss at both design and off-design working conditions, guaranteeing also the mechanical requirements. In the present work the authors, using a genetic algorithm (GA)-based tool, have explored the design space of modern axial flow compressor profiles reaching the target of better performance. The outlet flow angle and the mechanical-related quantities have been taken into account as design constraints. A fully integrated software procedure has been developed and applied to the redesign of existing airfoils in two different ways. The first redesign has been performed only at design condition while a further redesign has been implemented taking into account also two reference off-design conditions in order to increase the useful operating range. These reference off-design conditions are automatically obtained by using a loss curve model recently presented by the authors for prescribed velocity distribution airfoils. The computational tool has been applied for the optimization of the UKS-31 stator midspan section showing improved aerodynamic performance for both single and multi-point analysis.
Estimating Critical Flaw Sizes in Welds of the New ARES I-X Space Vehicle for Its Successful Solo Flight Using Probabilistic Approaches
Authors: Pai Shantaram S. , Hoge Peter A. , Patel B. M. , Nagpal Vinod K.
Published: August 2009

The primary structure of the ARES I-X Upper Stage Simulator (USS) launch vehicle is constructed of welded mild steel plates. There is some concern over the possibility of structural failure due to welding flaws. It was considered critical to quantify the impact of uncertainties in residual stress, material porosity, applied loads, and material and crack growth properties on the reliability of the welds during its pre-flight and flight. A criterion — an existing maximum size crack at the weld toe must be smaller than the maximum allowable flaw size — was established to estimate the reliability of the welds. A spectrum of maximum allowable flaw sizes was developed for different possible combinations of all of the above listed variables by performing probabilistic crack growth analyses using the ANSYS finite element analysis code in conjunction with the NASGRO crack growth code.
Computational Investigations Into Aero-Engine Bearing Chamber Off-Take Flows
Authors: Robinson Adam , Morvan Herve´ , Eastwick Carol
Published: August 2009

Within an aero-engine bearing chamber oil is provided to components to lubricate and cool. This oil must be efficiently removed (scavenged) from the chamber to ensure it does not overheat and degrade. Bearing chambers typically contain a sump section with an exit pipe leading to a scavenge pump. In this paper a simplified geometry of a sump section is analysed computationally. Two experimental rigs have been utilised to support this work, although these results are not reported in the present paper. Early data from the experimental work is both qualitative (visualisation), and quantitative (conductance probe measurements of film height, analysis of images); but this work is still ongoing. The results indicate that for the chosen geometry there are three separate regimes of flow dependent on inlet flow conditions. To complement the experimental studies the commercial computational fluid dynamics (CFD) code Fluent 6 has been applied to investigate typical flows found in aeroengine bearing chambers. Using the Volume of fluid (VOF) model of Hirt and Nichols, the CFD investigation has built from fundamental cases to computations for the experimental test rig configurations. This is the focus of the present paper. The first CFD case presented is a cylindrical vessel drained through a centrally placed cylindrical pipe. The geometry was chosen to match that of a published experimental study (Lubin and Springer [1]) and CFD data compares well with the available data. A full 3D model was required and a hexahedral mesh was implemented as it was found to provide a better discretisation. Results presented show that the volume of fluid approach is capable of modelling draining flows. Building on this, a three-dimensional CFD investigation into two phase flow in a double intake scavenge system similar to those found in aero-engines has been conducted. The paper shows that CFD is a promising approach to analysing bearing chamber scavenge flows and further, indicates methods of characterising behaviour.
Extending the Fuel Flexibility From Natural Gas to Low-LHV Fuel: Test Campaign on a Low-NOx Diffusion Flame Combustor
Authors: Giannini Nicola , Zucca Alessandro , Romano Christian , Ceccherini Gianni
Published: August 2009

Today’s Oil & Gas facility market requires enlarging machines’ fuel flexibility toward two main directions: on the one hand burning fuels with high percentages of Ethane, Butane and Propane, on the other hand burning very lean fuels with a high percentage of inerts. GE has extensive experience in burning a variety of gas fuels and blends in heavy-duty gas turbines. From a technical point of view, the tendency towards leaner fuel gases for feeding gas turbines, introduces potential risks related to combustion instability, on both combustion hardware and machines’ operability. GE Oil&Gas (Nuovo Pignone), has developed a new program aimed to extend the fuel flexibility of its Low-NOx diffusion flame combustor (Lean Head End, or LHE), which currently equips single and dual shaft 30 MW gas turbines, so that it can handle low-LHV fuels. A fuel flexibility test campaign was carried out at full and partial load conditions over an ambient and fuel range, in order to investigate both ignition limits and combustor performances, focusing on hot parts’ temperatures, pollutant emissions and combustion driven pressure oscillations. The pressurized tests were performed on a single combustion chamber, using a dedicated full-scale (full-pressure, full-temperature and full-flow) combustor test cell. Variable composition gaseous fuel mixtures, obtained by mixing natural gas with N2 from 0% up to about 50% vol., were tested. The experienced LHE combustion system up to now had been fed only with natural gas in multi can single gas combustion systems. Combustion system modifications and different burner configurations were considered to enlarge system capabilities, in order to accommodate operation on the previous mentioned range of fuel mixtures, including: nozzle orifice sizing and combustor liner modification. This paper aims to illustrate the upgraded technology and the results obtained. Reported data show combustion system’s performances, mainly in terms of pollutant emissions and operability. The performed test campaign demonstrated the system’s ability to operate at all required loads with diluted natural gases containing up to 50% vol. of N2. Results also indicate that ignition is possible with the same inerts concentration in the fuel, keeping the fuel flow at moderately low levels. As far as load operation, the combustion system proved to be almost insensitive to any tested inerts concentration, while a huge reduction of NOx emissions was observed increasing the molar fraction of N2 in the fuel gas, maintaining good flame stability.
Critical Components Life Update for Gas Turbine Engines: Case Study of an International Collaboration
Authors: Beres Wieslaw , Fread Donald , Harris Lesley , Haupt Philip , Kappas Joanna , Olson Roger , Reineke Philip , Robertson Sandi , Stocks Gordon
Published: August 2009

The paper describes results of the international collaboration that led to revision of the declared lives for critical components of a turbo-prop gas turbine engine. Four nations contributed to the program—Australia, Canada, USA and South Africa under the auspices of a Component Improvement Program led by the Original Equipment Manufacturer (OEM). This international collaboration was initiated as a result of the decrease in the declared life for some critical components of this engine by the OEM. The core of the program consisted of a detailed stress analysis performed in South Africa, and spin rig testing of selected life-limited, rotating turbine components—two stages of discs and two stages of spacers—performed in Australia and Canada. The general objectives of the program were to provide more accurate low cycle fatigue crack initiation data and to verify crack growth life analysis techniques using advanced 2D and 3D finite element analyses and spin rig testing for selected components. The crack initiation results are used to improve the life management procedures. Since the OEM does not recommend using life limits that exceed the safe crack initiation life of the rotating turbine components, the crack growth analysis results are used only for risk assessment and risk management by the engine operators. The basis of analytical techniques used for preparing the tests as well as the testing procedures are described. In addition, the development of NDE (Non Destructive Evaluation) methods and the inspections of these components during and after the tests are discussed. The economical benefits of such an international collaboration are demonstrated. The uniqueness of this approach to life revision of critical components of gas turbine engines, particularly for engines that have been in operation for many years, includes close cooperation of an international team of the engine manufacturer, the major engine users and their respective scientific organizations. In addition, a significant amount of operational experience that has been accumulated by the OEM, has allowed for verification of the spin rig test results.
A Versatile Design of a Controlled Swirl Distortion Generator for Testing Gas Turbine Engines
Authors: Sheoran Yogi , Bouldin Bruce
Published: August 2009

Inlet swirl distortion has recently become a major area of concern in the gas turbine engine community. Gas turbine engines are being installed in embedded installations that are downstream of increasingly complicated inlet systems, such as those used in Unmanned Aerial Vehicles (UAVs). These inlet systems can produce complex swirl patterns in addition to total pressure distortion. The effect of swirl distortion on engine or compressor performance and operability must be evaluated. The gas turbine community is developing methodologies to measure and characterize swirl distortion. There is a strong need to develop a mechanism for generating prescribed swirl distortion intensities and patterns for testing compression system sensitivity to swirl distortion. Several devices such as delta wings or complex turning vanes have been proposed and used to generate swirl distortion with limited success. This paper presents a versatile swirl distortion generator design that produces a wide range of swirl distortion patterns of a prescribed strength, including bulk swirl, 1/rev, and 2/rev patterns. It also creates different paired swirl patterns, varying from equal and opposite “twin swirls” to offset swirl pairs. This paper describes the design of the swirl generator. Computational Fluid Dynamics (CFD) results are presented along with some test data which illustrate how the swirl generator functions and how altering the swirl generator configuration can produce different swirl distortion patterns.
Three-Dimensional CFD Analysis of a Gas Turbine Combustor for Medium/Low Heating Value Syngas Fuel
Authors: Xiong Yan , Ji Lucheng , Zhang Zhedian , Wang Yue , Xiao Yunhan
Published: August 2009

Gas turbine is one of the key components for integrated gasification combined cycle (IGCC) system. Combustor of the gas turbine needs to burn medium/low heating value syngas produced by coal gasification. In order to save time and cost during the design and development of a gas turbine combustor for medium/low heating value syngas, computational fluid dynamics (CFD) offers a good mean. In this paper, 3D numerical simulations were carried out on a full scale multi-nozzle gas turbine combustor using commercial CFD software FLUENT. A 72 degrees sector was modeled to minimize the number of cells of the grid. For the fluid flow part, viscous Navier-Stokes equations were solved. The realizable k-ε turbulence model was adopted. Steady laminar flamelet model was used for the reacting system. The interaction between fluid turbulence and combustion chemistry was taken into account by the PDF (probability density function) model. The simulation was performed with two design schemes which are head cooling using film-cooling and impingement cooling. The details of the flow field and temperature distribution inside the two gas turbine combustors obtained could be cited as references for design and retrofit. Similarities were found between the predicted and experimental data of the transition duct exit temperature profile. There is much work yet to be done on modeling validation in the future.
A New Method for Measuring Thermal Contact Conductance: Experimental Technique and Results
Authors: Woodland Simon , Crocombe Andrew D. , Chew John W. , Mills Stephen J.
Published: August 2009

Thermal contact conductance (TCC) is used to characterise heat transfer across interfaces in contact. It is important in thermal modelling of turbomachinery components and finds many other applications in the aerospace, microelectronic, automotive and metal working industries. A new method for measuring TCC is described and demonstrated. A test rig is formed from an instrumented split tube with washers in-between and loading applied in controlled conditions. The experimental method and data analysis is described, and the effect on thermal contact conductance of parameters such as contact pressure, surface roughness, surface flatness and loading history is investigated. The results of these tests are compared to those in the available literature and good agreement of trends is found. However, the tests conducted to measure the effect of load cycling on TCC have found that the TCC continues to increase beyond 20 or so load cycles, contrary to some results in the literature.
Simulation of Turbulent Flows in Turbine Blade Passages and Disc Cavities
Authors: Volkov Konstantin N. , Hills Nicholas J. , Chew John W.
Published: August 2009

The main objectives of the paper are to test widely used turbulence models against selected benchmark problems identifying range of applicability and limitations of the models used and to help accurately predict flow in turbine blade passages and disc cavities. The following models are considered: the k–ε model with and without a Kato–Launder correction and a Richardson number correction, the two-layer k–ε/k–l model, and the Spalart–Allmaras model with and without correction for rotation. Weaknesses in the models are identified and suggestions made for possible improvements. Numerical implementation of the wall functions approach is also considered. The test cases considered include flat plate boundary layer, flat plate heat transfer, enclosed rotating disc, and a combined turbine blade/disc cavity model. Comparisons are made with experimental data and computations from different CFD codes.
Experimental and Numerical Investigations on Turbine Airfoil Cooling Designs: Part I—An Investigation on Flow Features by Particle Image Velocimetry
Authors: Wang J. H. , Xu H. Z. , Liu Y. L. , Du Z. N. , Yang S. J.
Published: August 2009

Experimental and numerical investigations are conducted to understand the features of the fluid dynamics within double-wall film-cooled configurations. Based on the similarity principle of the Reynolds number, a large-scale similar configuration made of transparent material is used as specimen, and the fluid velocity distributions over several typical cross sections within the specimen channel are captured by a particle image velocimetry (PIV) system. The experiments are carried out at a density ratio of fluid medium to tracer particle 1.05. The flow features are respectively calculated by different turbulence models and numerical grids. To confirm turbulence models and numerical grids, the numerical results are compared with the experimental data obtained by the PIV system. Through the comparisons, recommendations have been made with regard to the best model and numerical grid which best predict such velocity fields. The influences of inlet Reynolds numbers and the geometrical device of the double-wall film-cooled configurations on the features of flow field are numerically simulated by the recommended model and grid. The simulation results predicate that the flow features are mainly dominated by the geometrical device, the inlet Reynolds number can only result in a magnitude change of velocity fields, and this change is almost linear. This is the first part of the entire investigations on the double-wall film-cooled configurations, and the objective of this part is to confirm a suitable mathematical model and numerical grid for describing the flow features. In the next part, the overall heat transfer characteristics of these configurations will be studied.
Experimental and CFD Based Determination of Flutter Limits in Supersonic Space Turbines
Authors: Groth Pieter J. , Ma˚rtensson Hans E. , Edin Niklas
Published: August 2009

Turbines operating at high pressure in high velocity flow are susceptible to flutter. As reduced frequencies become sufficiently low, negative aerodynamic damping will be found in some modes. Ensuring that the total system damping is positive over the entire turbine operating envelope for all modes is of utmost importance in any design since flutter in a turbine often causes blade failures. This is in contrast to the normal engineering approach, which is to require a positive aerodynamic damping. A unique test campaign with a 1.5 stage supersonic space turbine has been performed. The turbine was operated at simulated running conditions over a large operating envelope in order to map out flutter limits. During the test, flutter was intentionally triggered at seven different operating conditions. Unique data have been obtained during the test that supports validation of design tools and enables better understanding of flutter in this type of turbine. Based on the data the flutter boundary for the turbine could be established. Using CFD tools flutter was predicted at all operating points where the flutter limit was crossed. Both in predictions and as evidenced in test the 2 nodal diameter backward traveling mode was the most unstable. In addition to this predicted values of aerodynamic damping at flutter agreed well with damping estimated from measured amplitude growth.
Design and Experimental Verification of Mistuning of a Supersonic Turbine Blisk
Authors: Groth Pieter , Ma˚rtensson Hans , Andersson Clas
Published: August 2009

A rotor blisk of a supersonic space turbine has previously been designed to allow for free flutter to occur in an air test rig [1]. Flutter occurred at several operating conditions and the flutter boundary for the test turbine was established. In this paper the rotor blisk is redesigned in order to inhibit flutter. The design strategy chosen is to introduce a mistuning concept. Based on aeroelastic analyses using a reduced order model (ROM) a criterion for the required level of mistuning is established in order to stabilize the lower system modes. Proposals in literature suggest and analyze mistuning by varying blade mode frequencies in random patterns or by modifying blades in an odd-even pattern. Here a modification of sectors of the blisk is introduced in order to bring a sufficient split of the system mode frequencies. To verify that the redesigned blisk efficiently could inhibit flutter an experiment similar to that in [1] is performed with the mistuned rotor blisk. By running the redesigned blisk at operating conditions deep in to the unstable region of the tuned blisk it is demonstrated that a relative low level of mistuning is sufficient to eliminate rotor flutter.
Aerothermal Performance of Streamwise and Compound Angled Pulsating Film Cooling Jets
Authors: Aga Vipluv , Mansour Michel , Abhari Reza S.
Published: August 2009

The quantification of aerothermal loss is carried out for streamwise and compound angled film cooling jets with and without large scale pulsation. This paper reports on the simultaneous measurements of the unsteady pressure and temperature field of streamwise and a 60° compound angled film cooling jet, both with a 30° surface angle over a flat plate with no pressure gradients. Turbine representative non-dimensionals in terms of the geometry and operating conditions are studied. The main flow is heated, more than the injected flow to have a temperature difference and hence a density ratio of 1.3, while the blowing ratio is maintained at 2. The entropy change, derived from pressure and temperature measurements, is calculated by using modified reference conditions to better reflect the losses in both the jet and the freestream. The effects of the periodic unsteadiness associated with rotating machinery are simulated by pulsating the jets. These effects are documented through time-resolved entropy change contours. Mass averaged entropy and kinetic energy loss coefficients seem to be apt quantities for comparing aerothermal performance of streamwise and compound angled injection. It is observed that the mass averaged entropy loss of a streamwise jet doubles when it is pulsated whereas that of a compound angled jet increases by around fifty percent. It may be conjectured from the measurements shown in this study that streamwise oriented jets suffer most of their entropy losses at the hole exit due to separation whereas in compound angled jets, downstream thermal mixing between the jet and the freestream, is the dominant mechanism.
Application of CFD to Assess the Performance of a Novel Pre-Swirl Configuration
Authors: Snowsill G. D. , Young C.
Published: August 2009

The urgent requirement to minimise the environmental impact and costs associated with gas turbine operation has provided the incentive for OEMs to design engines capable of performing at ever-increasing cycle pressure ratios, turbine inlet temperatures and rotational speeds. An inevitable consequence of this trend is the requirement for safe operation of gas turbine components in an increasingly hostile environment. The severity of the environment in which critical gas turbine components operate is such that even quite modest improvements in cooling effectiveness can result in a disproportionate improvement in component life. To this end, the technique of pre-swirling the cooling air delivered to rotating components, in order to reduce its relative total temperature and thereby increase cooling effectiveness, is well established [1]. The present work describes a novel pre-swirl arrangement which overcomes significant shortcomings found in more traditional systems. The benefits of the system and its sensitivity have been explored using CFD techniques. This investigation has shown that there are likely to be substantial benefits of up to 20K in relative total temperature from adopting such a system. The use of CFD to explore such a design allows considerable insight to be gained into its characteristics before investing heavily in experimental demonstration, on an engine or rig.
eXclean®: A New Surface Preparation of Turbine Components for Deposition of MCrAlY Coatings With Super Clean Interface
Authors: Russo Dario , Scrivani Andrea , Rizzi Gabriele , Lanzi Alessandro , Giolli Carlo
Published: August 2009

The most commonly used structural materials for blades and other high temperature components of gas turbines are nickel base superalloys. A TBC protection coating system consists of a top coat of yttria partially stabilized zirconia and an underlying bond coat, usually MCrAlY (where M stands for Ni, Co or a combination of both). MCrAlYs are normally deposited by the thermal spray processes: air plasma spray (APS), vacuum plasma spray (VPS/LPPS) or high velocity oxygen fuel (HVOF). The adhesion between the bond coat and the substrate, and therefore of the whole thermal barrier system, strongly depends upon the surface roughness. A high level of roughness generally denotes better adhesion, especially with the HVOF thermal spray process where it is a necessity. Generally the roughness is reached by means of grit blasting with an abrasive media; this results in a certain level of surface contamination due to the entrapment of abrasive particles. The aim of this work was to set up a new surface preparation process in order to obtain a completely clean surface with a suitable roughness, which can be coated afterwards with HVOF or VPS/LPPS thermal spray technology. The tests carried out by this process on turbine blades, coated with a HVOF system, led to obtaining a coating/base material interface without any contamination caused by the surface preparation.
A Non-Dimensional Quasi-3D Blade Design Approach With Respect to Aerodynamic Criteria
Authors: Dutta Amit Kumar , Flassig Peter Michael , Bestle Dieter
Published: August 2009

The competition between aero-engine manufacturers has increased dramatically in the last decades. Saving computational time within the design process, which is equivalent to saving money, is of major importance for the industry. Talking about the aerodynamic compressor blading process, it becomes indispensable to go for new or alternative ways in designing blades in order to fulfill raised performance demands. The focus of this paper, therefore, is to propose a quasi-3D aerodynamic design concept with extended and improved parameterization of the aerofoil in order to support the industrial blading process. A Be´zier-surface is selected to parameterize the non-dimensional camber-line angle distribution along the blade chord from leading to trailing edge over the entire blade height in radial direction. Starting from scratch, the geometric blade build-up is completed by superposing the resulting camber-line with a given thickness distribution. For additional increase of design freedom, Be´zier-curves are used to radially parameterize blade inlet and outlet angles in their dimensionless form. The chosen parameterization of these distributions guarantees smooth blade shapes and geometry distributions with a minimum of design parameters. For optimization purpose it is essential to get performance information on the entire blade, however, with minimal computational effort. Facing this challenge, aerodynamic blade performance is evaluated by a two-dimensional blade-to-blade flow solver for specific sections on different radial blade heights. In order to speed up the blade design process, the flow calculations are realized by a distributed computing concept on a Linux high-performance cluster. All investigations are carried out for highly loaded controlled diffusion blades which are taken from an existing industrial research application. Since selected criteria such as mean loss at design point conditions and working range for off-design flow conditions represent contradicting design goals, the blade design problem is solved by means of a multi-objective problem formulation and a stochastic optimization algorithm. As a result Pareto-optimal trade-off solutions between conflicting design goals are shown where the design engineer can choose from according to his specific preferences.
Combustion Simulation of an EGR Operated Micro-Gas Turbine
Authors: Cameretti Maria Cristina , Piazzesi Renzo , Reale Fabrizio , Tuccillo Raffaele
Published: August 2009

Following their recent experiences in the search of methods for reducing the nitric oxide emissions from a micro-gas turbine, the authors discuss in this paper the results of the combustion simulation under different conditions induced by the activation of an exhaust recirculation system. The theoretical approach starts with a matching analysis of the EGR equipped micro-turbine, and then proceeds with the CFD analysis of the combustor. Different combustion models are compared in order to validate the method for NOx reduction by the point of view of a correct development of the chemically reacting process.
Aero-Thermal Investigation of a Rib-Roughened Trailing Edge Channel With Crossing-Jets: Part II—Heat Transfer Analysis
Authors: Coletti Filippo , Armellini Alessandro , Arts Tony , Scholtes Christophe
Published: August 2009

The present contribution addresses the aero-thermal experimental and computational study of a trapezoidal cross-section model simulating a trailing edge cooling cavity with one rib-roughened wall and slots along two opposite walls. Highly resolved heat transfer distributions for the geometry with and without ribs are achieved using a steady state liquid crystals method in part II of this paper. The reference Reynolds number, defined at the entrance of the test section, is set at 67500 for all the experiments. Comparisons are made with the flow field visualizations presented in part I of the paper. The results show the dramatic impact of the flow structures on the local and global heat transfer coefficients along the cavity walls. Of particular importance is the jet deflected by the rib-roughened wall and impinging on the opposite smooth wall. The experimental results are compared with the numerical predictions obtained using the finite volume, Reynolds-Averaged Navier-Stokes solver CEDRE.
Experimental and Numerical Flutter Investigation of the 1st Stage Rotor in 4-Stage High Speed Compressor
Authors: Johann E. , Mu¨ck B. , Nipkau J.
Published: August 2009

Experimental tests were performed to investigate flutter behaviour of the transonic rotor in a high-speed multistage compressor test facility. Besides the acquisition of overall performance parameters the rig was equipped with special instrumentation such as strain gauges, tip-timing system and dynamic pressure transducers. The 4-stage compressor comprises 3 variable vanes. The instrumentation was able to measure stall and flutter and forced responses of the rotor blades. The experimental data was used to validate the in-house aeroelastic solver. During testing flutter was triggered intentionally at part speed conditions with malscheduled variable vanes. The malschedule changed the flow incidences for the relevant rotor and induced flutter. An aeroelastic simulation was set up according to the flutter conditions found during the experiment. The measured boundary conditions were used to set up the aeroelastic simulation. The calculation shows the torsion mode with a negative aerodynamic damping which confirms the test results. The only difference between test and prediction is the nodal diameter, at which the flutter occurs.
Turbine Blade Tip Leakage Flow Control: Thick/Thin Blade Effects and Separation Line Studies
Authors: Stephens Julia E. , Corke Thomas , Morris Scott C.
Published: August 2009

An experiment was conducted in a linear cascade of Pratt and Whitney Pak-B turbine blades for an exit Mach number of 0.3 to simulate the flow in the tip-gap region of a low pressure turbine blade row. The experiment focused on the independent effects of thickness-to-gap (t/g) and gap-to-chord (g/c) ratios on the tip-gap flow behavior. Two extreme g/c ratios of 5% and 8% were chosen, for which four tip t/g ratios were simulated using pressure-side winglets. The flow was documented through blade-tip and end-wall static pressure measurements, and downstream total pressure loss coefficients. Additionally, surface flow visualization was performed on the blade tip end for a greater understanding of the gap-flow behavior. The response of the flow to passive flow control using a partial suction-side squealer tip at each of the t/g and g/c cases was documented. The intention was to examine any sensitivity of the flow to the g/c ratio that might be attributed to the t/g ratio in a manner that can be categorized as “thick” or “thin” blade behavior. For this, the focus was on possible changes in the size and location of separation and reattachment lines on the blade tip end. The results presented in this paper indicate that the behavior of the flow in the tip-gap region of a linear cascade turbine blade depend both on t/g and the g/c ratios. Downstream loss coefficients generally decreased with increasing t/g for the 5% g/c case, while they were relatively steady with a slight increase with increasing t/g for the 8% g/c case. The squealer reduced pressure loss coefficient for both g/c cases. It was seen to have a peak effectiveness at t/g of about 3.7 for the 5% g/c case, and at t/g of 3.3 for the 8% g/c case, with diminishing effectiveness at larger t/g ratios. Blade loading showed similar dependence on both g/c and t/g. For the baseline flat-tip case, the tip loading initially decreased with increasing t/g up to about 3.5. For values greater than 3.5, the tip loading increased slightly. The addition of the squealer tip increased the tip loading for all of the t/g ratios at the 5% g/c ratio. The trend with t/g generally followed that of the baseline flat-tip case. However, for the 8% g/c case, the blade loading increased almost linearly with increasing t/g, and the squealer increased loading at the smallest t/g but decreased loading for all other t/g cases investigated. In all cases regardless of the t/g, the surface flow visualization revealed a well defined separation and reattachment region. However, the chordwise location of the start of the separation moved toward the trailing edge with increasing t/g and, independently, with increasing g/c. Therefore, no “thick” or “thin” blade regimes were found, and flow characteristics were determined to depend on both g/c and t/g ratios.
An Analytical Approach for Gas Turbine Parameter Corrections
Authors: Kyritsis Vasileios E. , Pilidis Pericles
Published: August 2009

Turbomachinery component behavior depends on dimensionless parameters, such as inlet and circumferential Mach numbers and the ratio of specific heats. Regarding mass flow and speed, their dimensionless scaling parameters are usually used instead based on Mach number similarity. A given dimensionless aerodynamic operating point is defined by certain values of axial and circumferential Mach numbers. To such a point and for a certain value of isentropic exponent, a given dimensionless enthalpy variation corresponds as the work parameter. When turbo-machinery performance sizes, such as the work parameter and efficiency, are plotted against mass flow and speed to form a characteristic, the absence of the isentropic exponent as an additional dimension causes inaccuracies. The extent of the inaccuracies firstly depends on the scaling groups used for mass low, speed and work, that is whether they include the gas property parameters, such as the isentropic exponent and the gas constant. The aforementioned shows that for rigorous calculations correction factors have to be applied, especially when quasi-dimensionless groups are used and/or pressure ratio is used as the work parameter. Typically, the corrections for mass flow and speed may take the form of multipliers, which consist of ratios of the isentropic exponent and/or the gas constant between the examined condition and the reference one. Alternatively, for the case of mass flow the exponents of temperature and pressure can deviate from their theoretical values of 0.5 and 1.0 respectively. Scope of the current work is the mathematical formulation of such exponents for a variety of scaling groups regarding mass flow, speed and work. The correction factors are a strong function of the operating condition, temperature and gas composition, which fully define gas properties. Among the findings of the current study, evidence is provided that the practically one-to-one relationship considered between dimensionless mass flow and inlet Mach number holds for low Mach number values. This is particularly true, since its sensitivity to variations of the isentropic exponent gets increasingly larger with Mach number. Additionally, for a given dimensionless enthalpy change, the exchange rate of pressure ratio against variations of the isentropic exponent is much more increased for an expansion rather than a compression. The latter justifies the use of dimensionless enthalpy drop in turbine characteristics.
The Effects of Aerodynamic Asymmetric Perturbations on Forced Response of Bladed Disks
Authors: Miyakozawa Tomokazu , Kielb Robert E. , Hall Kenneth C.
Published: August 2009

Most of the existing mistuning research assumes that the aerodynamic forces on each of the blades are identical except for an interblade phase angle shift. In reality, blades also undergo asymmetric steady and unsteady aerodynamic forces due to manufacturing variations, blending, mis-staggered blades or in-service wear or damage, which cause aerodynamically asymmetric systems. This paper presents the results of sensitivity studies on forced response due to aerodynamic asymmetry perturbations. The focus is only on the asymmetries due to blade motions. Hence, no asymmetric forcing functions are considered. Aerodynamic coupling due to blade motions in the equation of motion is represented using the single family of modes approach. The unsteady aerodynamic forces are computed using CFD methods assuming aerodynamic symmetry. Then, the aerodynamic asymmetry is applied by perturbing the influence coefficient matrix in the physical coordinates such that the matrix is no longer circulant. Therefore, the resulting aerodynamic modal forces in the traveling wave coordinates become a full matrix. These aerodynamic perturbations influence both stiffness and damping while traditional frequency mistuning analysis only perturbs the stiffness. It was found that maximum blade amplitudes are significantly influenced by the perturbation of the imaginary part (damping) of unsteady aerodynamic modal forces. This is contrary to blade frequency mistuning where the stiffness perturbation dominates.
Calibration of CFD Model for Mist/Steam Impinging Jets Cooling
Authors: Wang Ting , Dhanasekaran T. S.
Published: August 2009

In the heavy-frame advanced turbine systems, steam is used as a coolant for turbine blade cooling. The concept of injecting mist into the impinging jets of steam was experimentally proved as an effective way of significantly enhancing the cooling effectiveness in the laboratory under low pressure and temperature conditions. However, whether mist/steam cooling is applicable under actual gas turbine operating conditions is still subject to further verification. Recognizing the difficulties of conducting experiments in an actual high-pressure, high-temperature working gas turbine, a simulation using a CFD model calibrated with laboratory data would be an opted approach. To this end, the present study conducts a CFD model calibration against the database of two experimental cases including a slot impinging jet and three rows of staggered impinging jets. Using the experimental results, the CFD model has been tuned by employing different turbulence models, computational cells, wall y+ values, and selection of near-wall functions. In addition, the effect of different forces (e.g. drag, thermophoretic, Brownian, and Saffman’s lift force) are also studied. None of the models are good predictors for all the flow regions from near the stagnation region to far-field downstream of the jets. Overall speaking, both the standard k-ε and RSM turbulence models perform better than other models. The RSM model has produced the closest results to the experimental data due to its capability of modeling the non-isotropic turbulence shear stresses in the 3-D impinging jet fields. For the 3-D flow fields, the nearest element from the wall must be set to approximately unity (y+ ≈ 1) to capture the correct flow structure. The simulated results showed that the calibrated CFD model could predict the heat transfer coefficient of steam-only case within 2 to 5% deviations from the experimental results for all the cases. When mist is employed, the prediction of wall temperatures is within 5% for a slot jet and within 10% for three-row jets.
Experimental Analysis of Varied Vortex Reducer Configurations for the Internal Air System of Jet Engine Gas Turbines
Authors: Gu¨nther Andre´ , Uffrecht Wieland , Kaiser Erwin , Odenbach Stefan , Heller Lothar
Published: August 2009

The continuous improvement of engine performance, combined with strict environmental and safety regulations and the reduction of time and cost of new products, is the major goal of the turbomachinery industry. Particular attention is being focused on the reduction of internal losses and weight, associated with the internal air system. The cooling air is normally bled by holes in the rotor from the main flow of the HP compressor, transported radially inwards towards the shaft and further transferred to the hot parts of the engine. The radial inflow creates vortices induced by the core rotation ratio, which create very high pressure losses and restrict the maximum cooling flow rate. The pressure loss depends strongly on the rotating speed and the mass flow rate. The vortex reducer prevents the development of vortices and therefore reduces the pressure loss. A key area of concern is to optimize the pressure loss concurrent with the use of new light weight or easy to manufacture configurations of vortex reducers. The material presented in this paper describes an experimental study, concentrating on a two cavity test rig for different internal cooling flow concepts. The test rig has steel discs, operating at engine representative flow and temperature conditions and permits several flow and heating modes with axial or/and radial flow configurations. The present work investigates the fluid flow for different vortex reducer configurations at different rotational speeds and its influence on the pressure loss. Particular attention was paid to the influence of size and location of the tubes. The experimental setup and the results concerning the pressure losses for the different configurations are presented.
Validation of CFD Predicted Discharge Coefficients for Thick Plate Orifices With Approach Flow Perpendicular and Inclined to the Orifice Axis
Authors: Gupta Anish Kumar , Ramerth Douglas , Ramachandran Dhinagaran
Published: August 2009

The flow discharge through thick orifices with approaching flow normal and inclined to the orifice axis was numerically predicted. The objective is to validate Fluent code predictions of discharge coefficient against Rohde et al. experimental data and arrive at an optimum mesh in terms of effort, simulation time and accuracy. CFD simulations were performed on several Rohde et al. experimental models with orifice thickness to length ratio varying from 0.5 to 4.0, with sharp and rounded inlet edges and orifice axis inclined 45 degrees to the approaching flow. The approach Mach number varies from 0.07 to 0.55 and orifice velocity head ratio ranges from 1.3 to 200. Simulations are performed using FLUENT V6.3 with k-ω SST model and 2nd order discretization scheme. Coarse mesh results at selected test points were compared with fine mesh results and a best meshing practice was determined. CFD computation at selected points was also performed with Realizable K-ε, v2f and RSM turbulence model. CFD predictions of discharge coefficient show good agreement with Rohde et al. experimental data.
Dynamic and Flashback Characteristics of the Syngas Premixed Swirling Combustors
Authors: Song Quanbin , Fang Aibin , Xu Gang , Xu Yanji , Huang Weiguang
Published: August 2009

Flashback and combustion thermoacoustic oscillation are two of the major problems in the lean premixed combustors of gas turbine, especially for hydrogen enriched syngas fuels. This paper introduced the experimental results regarding the characteristics research of thermoacoustic oscillation, dynamic heat release and flashback based on a premixed combustion experimental system. The critical flashback equivalence ratio and three dynamic parameters were tested in four types of combustor using eleven kinds of simulated syngas which have the same LHV. Partial flashback and full flashback are observed. With the increase of the hydrogen concentration, the critical equivalence ratio is decreasing and full flashback occurs more easily. The change of nozzle’s geometry structure has a consequent effect on the flashback characteristics. Flashback occurs more easily when the thermoacoustic oscillation is greater, and the corresponding critical equivalence ratio is smaller. Both flashback status and normal status are relative stable combustion within a certain range of equivalence ratio. As the generation and disappearance of the thermoacoustic oscillation and its amplitude-frequency characteristics somehow could be in some randomicity, the critical equivalence ratio is correspondingly not a constant. The dynamic chemiluminescence intensity which indicates the dynamic heat release has a good corresponding relationship in terms of amplitude and frequency characteristics with the dynamic pressure which indicates the thermoacoustic oscillation. Through power spectral density analysis, dynamic pressure and chemiluminescence intensity signals both reflect the dynamic characteristics of the combustor and can be applied in the research of the combustion instability, including the thermoacoustic oscillation characteristics, fluctuation of the chemical heat release rate and flashback characteristics and so on. The results and methods could give some reference to the development of hydrogen enriched gas turbines for future IGCC and poly-generation systems.
Influence of Compressor Performance Maps Shape on Wet Compression
Authors: Bhargava Rakesh K. , Bianchi Michele , Melino Francesco , Peretto Antonio , Spina Pier Ruggero
Published: August 2009

In recent years, a great number of studies were carried out in order to analyze the main features of fogging technologies. The various fogging strategies seem to improve gas turbine and combined cycle power output with low initial investment cost and less installation downtime. In fact, nowadays fogging is successfully installed on several gasturbine and combined cycle power plants worldwide. In particular, overspray fogging and interstage injection involve two-phase flow consideration and water evaporation during compression process (also known as wet compression). The aim of the present paper is to further improve understanding of the wet compression process including stage-by-stage compressor behavior by investigating the influence of the axial compressor performance map shape on the evaporation process of the injected water through the compressor, achievable power boost, the maximum amount of water which can be injected and/or influence on the surge conditions. This analysis is carried out by using a calculation code, named IN.FO.G.T.E. (INterstage FOgging Gas Turbine Evaluation), developed and validated by the Authors.