bluuebunny/crossref_metadata_2025 huggingface dataset:
Impact of Flowfield Unsteadiness on Film Cooling of a High Pressure Turbine Blade
Authors: Vitt Paul , Iverson Chad , Malak Malak F. , Liu Jong S.
Published: January 2011

A key factor in gas turbine blade operation is flow unsteadiness generated by the vane as the blade passes behind it, and the impact of this variation on the aerodynamic performance of the blade has been well studied. The objective of the current study was to examine the impact of this flow unsteadiness on film cooling effectiveness on the airfoil. A Honeywell research single-stage high-pressure cooled turbine was selected as an engine-representative geometry. Both steady-state and transient stage calculations were completed using fully cooled vane and blade models. One area of interest was the impact of the vane passing on the local film hole flow rates, so the blade model included the leading edge impingement chamber as well as detailed film cooling holes in the analysis. Results from the calculations showed a dramatic change in the predicted film cooling on the blade surface in two primary areas. At the leading edge, the steady-state model predicted low local film effectiveness due to the stagnation line falling between showerhead film rows, whereas the time-averaged unsteady analysis showed an even distribution of film at the leading edge. On the pressure side, the steady model predicted effective film coverage downstream of the gill holes. In the unsteady analysis, the low momentum mixed gas and film stream along the forward portion of the pressure side was highly disturbed by the vane unsteadiness, and time-averaging the results indicated a significant reduction in overall film cooling effectiveness. The results show that considering the unsteady operating environment of the blade is very important during design and optimization of the blade thermal protection system.
Investigation of Augmented Angel Wing Sealing Through the Use of Surface Concavity Arrays
Authors: Bunker Ronald S. , Sheng Nuo
Published: January 2011

A simple rotor-stator disk experiment is used to investigate the enhancement in rim sealing effectiveness due to a surface treatment applied to the angel wings on either the rotor or stator. Angel wings in rim seals are also commonly referred to as seal teeth or flow discouragers. In the present study, arrays of hemispherical concavities, also known as dimples, have been placed in the surfaces of a single overlapping angel wing set at the outer radius of the rotor-stator system. Stator surface static pressure measurements are used to quantify the degree of external air ingestion into the disk cavity. Tests were run at a single rotational Reynolds number of 1.45e6 and 4500 rpm. The nominal radial clearance gap was 1.5 mm. Three axial engagement positions of the angel wings were tested. The non-dimensional purge flow rate Cw ranged from 3500 to 10000. Addition of a concavity array surface treatment on the rotor angel wing had negligible effect on the sealing capability for the rotor-stator seal system. However, test results indicate an increase in seal effectiveness of nearly 20% due to the presence of the concavity array surface on the stator angel wing. Through other known concavity flow physics, it is conjectured that the ejected organized vortical flows from the concavities create an effective fluid resistance in the radial seal gap that decreases the ingestion flow rate.
A Model for Cylindrical Hole Film Cooling: Part II—Model Formulation, Implementation and Results
Authors: auf dem Kampe Tilman , Vo¨lker Stefan
Published: January 2011

A model to simulate flows ejected from cylindrical film cooling holes in 3D-CFD without meshing the cooling hole geometry has been developed. It uses a correlation-based prediction of the complete three-dimensional flow field in the vicinity of a film hole exit based on characteristic film cooling parameters that is presented in part I of this two-part paper [1]. The model describes the film-jet in terms of its shape and the distribution of temperature and velocity components within the film-jet body. For example, the characteristic counter-rotating vortex pair in the film-jet is modeled. Adding source terms to the transport equations for mass, momentum and energy locally, the correlation-based prediction of the film-jet flow field is imposed onto a 3D-CFD simulation. Source terms are specified in the vicinity of a film hole exit, within a region representative of the volume occupied by the film jet. Each node within this source volume is treated individually in order to model the complex flow structure of the film-jet. The model has successfully been implemented in a commercial CFD code. Its general applicability has been tested and proven. The model’s predictive capability is compared to detailed CFD calculations and experimental investigations. A grid requirement study has been conducted, showing that the film cooling model delivers reasonable predictions of the surface temperature distributions downstream of the ejection location using relatively coarse grids. A minimum grid resolution requirement has been identified.
Multidisciplinary 3D-Optimization of a Fan Stage Performance Map With Consideration of the Static and Dynamic Rotor Mechanics
Authors: Siller Ulrich , Aulich Marcel
Published: January 2011

Achievement of an optimal compressor design with respect to its aerodynamic performance and feasible structural mechanics within an automated optimization process is subject of this paper. The compressor considered is a highly loaded, transonic fan stage, designed for achievement of a very high pressure ratio. To ensure operation in highly integrated installation conditions, a sufficient stability margin is of major concern. Multiple aerodynamic operating points at two rotational speeds allowed optimization of both the stability margin and the working line stage efficiency. On the part of structural mechanics, several static stress criteria were addressed for definite blade regions as well as the dynamic blade behavior in terms of the Campbell diagram. An optimization strategy was chosen, which targeted firstly on the fulfillment of multiple mechanical and aerodynamical constraints, while the aerodynamic performance was under constraint itself. Upon achievement, optimization aimed for maximum aerodynamic performance while keeping mechanics feasible. Response surfaces have been incorporated in the optimization process to reconcile costly high fidelity CFD and structural simulations with the large number of 114 free design parameters. Furthermore, optimization on these models enabled a successfully accomplishment of the constraint issue by a large number of numerically cheaper fitness evaluations. Starting from an already optimized baseline configuration, the current work targeted an improvement of the rotor aerodynamics in the transonic hub region and the resolution of previously unsolved problems concerning the rotor structural mechanics. Free design parameters were hub and casing contours in the rotor part, the shape of the leading and trailing blade edges and a high degree of freedom for rotor profile sections in the lower half of the blade.
Euler Inverse Axisymmetric Solution for Design of Axial Flow Multistage Turbomachinery
Authors: Rosa Taddei Simone , Larocca Francesco , Bertini Francesco
Published: January 2011

An axisymmetric model is used that replaces the blades with hub-to-tip streamsurfaces and their main effects on the flow with blade force terms. Euler equations are solved by means of an upwind finite-volume scheme for a specified swirl distribution in the meridional blade regions. This quantity drives the spanwise distribution of specific work for rotors and the swirl adding/removing property for stators. The time-marching procedure includes an evolutionary equation for the hub-to-tip geometry in each blade region. This equation involves use of a specified leading edge as a blade stacking line. Throughout the computation, the streamsurfaces take that shape providing the desired swirl distribution. The method is suitable to highly loaded blades i.e. turbine design. Compared to a traditional through-flow method, it predicts the overall turbine performances with about a 1% error. A satisfactory agreement is found concerning the spanwise distribution of the core flow quantities as well.
Spreadsheet as a Data Acquisition Tool in Turbomachinery
Authors: Engeda Abraham
Published: January 2011

This paper shows the power of spreadsheets as a strong tool in engineering teaching and research labs. In applied thermo-fluid education, even the one dimensional design or simple experimental measurement and analysis becomes a very complex exercise unless the procedure is programmed. Due to lengthy calculations and iterations, simple solutions are not possible. Exercises have therefore been limited in the classroom. But recent advances in powerful spreadsheets have opened a simple and fast way of performing design and advanced measurements. In recent times due to the introduction of a variety of mathematical soft wares, students have been relived from unnecessary time consuming chores; and therefore, complex measurements can now be carried out more comprehensively and easily. This paper reports on an experimental investigation to determine the effect of the vaneless diffuser width on the unsteady flow performance of a centrifugal compressor stage, where the whole data processing was carried out using a spreadsheet both for the steady and unsteady characteristics. Two compressor configurations with different vaneless diffuser width were investigated at four different impeller speeds and compared in the frequency and time domain. Only one diffuser rotating stall but different types of impeller rotating stalls were detected. The experiments show that the diffuser has a strong influence on the flow in the impeller including in areas way upstream. Analysis of the results indicated: • With increasing diffuser width the onset of impeller rotating stall was shifted to lower flow rates. • With increasing diffuser width the frequencies of the rotating stalls decreased. • There is a common tendency in most of the experiments to lower numbers of rotating cells with increasing relative speed. The whole data acquisition, processing and presentation are carried out using Excel.
OH* Chemiluminescence and OH-PLIF Measurements in a Micro Gas Turbine Combustor
Authors: Hohloch Martina , Sadanandan Rajesh , Widenhorn Axel , Meier Wolfgang , Aigner Manfred
Published: January 2011

In this work the combustion behavior of the Turbec T100 natural gas/air combustor was analyzed experimentally. For the visualization of the flame structures at various stationary load points OH* chemiluminescence and OH-PLIF measurements were performed in a micro gas turbine test rig equipped with an optically accessible combustion chamber. The OH* chemiluminescence measurements are used to get an impression of the shape and the location of the heat release zones. In addition the OH-PLIF measurements enabled spatially and temporarily resolved information of the reaction zones. Depending on the load point the shape of the flame was seen to vary from cylindrical to conical. With increasing thermal power load the maximum heat release zones shift to a lifted flame. Moreover, the effect of the optically accessible combustion chamber on the performance of the micro gas turbine is evaluated.
Fracture Analysis of a First Stage Turbine Blade
Authors: Ghorbani Reza , Asadikouhanjani Saeed , Kusterer Karsten , Haj Ayed Anis
Published: January 2011

Blade failures in gas turbine engines often lead to the loss of all downstream stages and it can have a dramatic effect on the availability of the turbine engines. This paper presents the analysis of an in service failure of a first stage gas turbine blade. The premature failure of the blade, made of nickel-base superalloy Inconel 738 LC, occurred after a service life of 8,127 EOH with normal start/stop and caused extensive damage to the unit. Crack growth mechanism has been evaluated based on macroscopic and microscopic observations of the fracture surfaces. Chemical analyses were carried out to identify the possible causes of the failures by examining anomalies in the chemical composition and microstructure analysis through SEM observations. The analysis of the different regions of fracture surface shows that crack propagation is mainly related to fatigue mechanism. Typical fatigue striations could be identified under a homogeneous oxide layer. The crack propagation occurred in the pressure-suction side direction and the initial crack origin is located on the missing part near leading edge area. The impact marks on the first stage leading edge of the blade and the general damage of the turbine give indication that the crack initiation was caused by an impact of a broken piece from first stage vanes or another object of unknown source.
Redefining Propulsion and Power Systems for the United States Navy’s 21st Century Destroyer: DDG 1000
Authors: Bigley Andrew , Driscoll Matthew
Published: January 2011

For the past 40 years, the United States Navy has utilized a standardized machinery configuration on its surface combatant cruisers and destroyers. Large gas turbines (18.5 MW) directly coupled to a twin screw drive train and smaller gas turbine engines (2.5–3.0 MW) feeding a common electrical bus provided ships propulsion and power requirements. This consistent design approach afforded an opportunity for the Navy to hone its operational and maintenance strategies with a focus on enhancing reliability. DDG 1000 provides a unique machinery arrangement with which the Unites States Navy has minimal operational experience, with small and large Gas Turbine prime movers all producing power to an integrated power distribution network servicing both propulsion and ships service power requirements. This new all electric platform design produces some unique challenges for both the prime movers and electrical distribution. This paper explores gas turbine operating profile, reliability centered maintenance, transient engine response, power quality requirements and power distribution architecture as they apply to this new surface combatant. Comparisons will be drawn between the Navy’s legacy system applications with an emphasis on how the new ship design requires innovative support approaches. Additionally, contrasts are articulated between defined military specifications and testing requirements for legacy applications and the amorphous standards for dual spool applications.
Heat Transfer Performance of Fan-Shaped Film Cooling Holes: Part II—Numerical Analysis
Authors: Bianchini C. , Facchini B. , Mangani L. , Maritano M.
Published: January 2011

Fan-shaped holes are widely used to provide better cooling performances than cylindrical holes over a large range of different operating conditions. Main advantages of such solution include a reduced amount of cooling air for the same performance, increased part lifetime and fewer required holes. As the overall cooling performance of such holes is strictly related to the adiabatic effectiveness and heat transfer coefficient (HTC) increase due to the coolant injection, both issues should be investigated. A numerical analysis has been conducted on a laidback fan-shaped film cooling hole onto a flat plate with the aim of investigating the increase of heat transfer. A steady-state RANS analysis was performed at two different blowing ratios (1.25 and 2.5) with imposed heat flux on the heated wall reproducing the same conditions as in the experimental tests presented in the companion paper. Despite no temperature difference was imposed between main gas and coolant flow, adiabatic effectiveness maps were extracted from tracing distribution over the plate. Performances of four different eddy viscosity turbulence models have been tested: the Two-Layer model by Rodi both in the isotropic original formulation and with an anisotropic algebraic correction based on DNS data fitting as firstly proposed by Lakheal, the k–ω SST by Menter and the ν2–f by Durbin. All calculations were conducted with a 3D unstructured pressure-based compressible solver based on the open-source OpenFOAM® CFD platform. A detailed analysis of both the predicted flow field and thermal distribution in the domain was presented. The obtained results were compared with the experimental measurements showed in the companion paper both in terms of wall heat transfer coefficient and adiabatic effectiveness.
An Investigation Into the Effects of Highly Transient Flight Maneuvers With Heat and Mass Transfer on the T-38 Air Force Trainer Inlet
Authors: Hale Alan , Hughes Andrew , Sirbaugh Jim , Kidman David S.
Published: January 2011

The T-38 Talon currently serves as the primary United States Air Force (USAF) trainer for fighter aircraft. This supersonic trainer was developed in the 1960s but continues to be used today as the result of various modernization programs throughout its service life. The latest Propulsion Modernization Program (PMP) focused on improved takeoff performance of the T-38’s inlets, improved reliability of the twin J85 afterburning turbojet engines, and reduced drag with an improved exhaust nozzle design. The T-38’s inlet includes bleed holes upstream of the engine face to provide cooling airflow from the inlet to the engine bay. However, at various flight conditions, the bay air is pressurized relative to the inlet resulting in reverse flow of hot engine bay air into the inlet. This reverse flow causes total temperature distortion that may reduce the engine stability margin. Partial inlet instrumentation of the left engine was used to estimate the total temperature distortion associated with reverse flow; however, flight testing of highly transient maneuvers revealed levels of total temperature distortion greater than that predicted for reverse flow alone. This discovery led to the hypothesis that thermal energy storage of the aluminum inlet during transient flight maneuvers resulted in increased temperature distortion at the engine face. Flight data analysis demonstrated the need for a near-real-time thermal inlet distortion analysis capability. A two-dimensional (2D) transient axisymmetric heat and mass transfer model was developed through the use of a lumped-parameter boundary-layer model to simulate the inlet flow and determine the time-dependent inlet duct heat transfer. This model was validated with transient 2D computational fluid dynamics (CFD) and two flight maneuvers. The analysis of flight maneuvers revealed that in the absence of engine bay air re-ingestion the time lag associated with the heating and cooling of the inlet walls generates radial temperature distortion which has the effect of reducing engine stability margin up to 5.44% for the maneuvers analyzed.
Effects of an Upstream Cavity on the Secondary Flow in a Transonic Turbine Cascade
Authors: Abo El Ella H. M. , Sjolander S. A. , Praisner T. J.
Published: January 2011

This paper examines experimentally the effects of an upstream cavity on the flow structures and secondary losses in a transonic linear turbine cascade. The cavity approximates the endwall geometry resulting from the platform overlap at the interface between stationary and rotating turbine blade rows. Previous investigations of the effects of upstream cavity geometries have been conducted mainly at low-speed conditions. The present work aims to extend such research into the transonic regime with a more engine representative upstream platform geometry. The investigations were carried out in a blow-down type wind tunnel. The cavity is located at 30% of axial-chord from the leading edge, extends 17% of axial-chord in depth, and is followed by a smooth ramp to return the endwall to its nominal height. Two cascades are examined for the same blade geometry: the baseline cascade with a flat endwall and the cascade with the cavity endwall. Measurements were made at the design incidence and the outlet design Mach number of 0.80. At this condition, the Reynolds number based on outlet velocity is about 600,000. Off-design outlet Mach numbers of 0.69, and 0.89 were also investigated. Flowfield measurements were carried out at 40% axial-chord downstream of the trailing edge, using a seven-hole pressure probe, to quantify losses and identify the flow structures. Additionally, surface flow visualization using an ultra-violet reactive dye was employed at the design Mach number, on the endwall and blade surfaces, to help in the interpretation of the flow physics. The experimental results also include blade-loading distributions, and the probe measurements were processed to obtain total-pressure loss coefficients, and stream-wise vorticity distributions. It was found that the presence of the upstream cavity noticeably altered the structure and the strength of the secondary flow. Some effect on the secondary losses was also evident, with the cavity having a larger effect at the higher Mach number.
Optimization of High Speed Centrifugal Compressor for a Micro Gas Turbine Based on CFD and FEM Analysis
Authors: Wang Hong-liang , Xi Guang
Published: January 2011

The centrifugal compressor impeller plays a significant role in the micro gas turbine. Its aerodynamic performance and structural reliability has great impact on the gas turbine system. Its aerodynamic performance and structural reliability are determined by blade design parameters and impeller back face design parameters. This paper presents a process of optimizing a centrifugal compressor impeller during the development of 100kW class micro gas turbine. Predicted by a 3D Navier-Stokes solver the aerodynamic performance is mainly affected by meridian contour and blade angle β distribution. The stress is calculated by means of a finite element analysis method and controlled by blade angle β distribution and thickness distribution at the hub. The requirement of high efficiency is met by using genetic algorithm and artificial neural network. Then the impeller back face profile is parameterized and optimized to decrease the stress of the impeller. The isentropic efficiency of the compressor is increased by 1% and the blade stress is reduced by 10% as compared with the original design. The results show that it is possible to reduce the centrifugal stress in the impeller without penalizing the aerodynamic performance through the optimization method.
Performance Study of a 1 MW Gas Turbine Using Variable Geometry Compressor and Turbine Blade Cooling
Authors: Bringhenti Cleverson , Tomita Jesuino Takachi , Barbosa Joa˜o Roberto
Published: January 2011

This work presents the performance study of a 1 MW gas turbine including the effects of blade cooling and compressor variable geometry. The axial flow compressor, with Variable Inlet Guide Vane (VIGV), was designed for this application and its performance maps synthesized using own high technological contents computer programs. The performance study was performed using a specially developed computer program, which is able to numerically simulate gas turbine engines performance with high confidence, in all possible operating conditions. The effects of turbine blades cooling were calculated for different turbine inlet temperatures (TIT) and the influence of the amount of compressor-bled cooling air was studied, aiming at efficiency maximization, for a specified blade life and cooling technology. Details of compressor maps generation, cycle analysis and blade cooling are discussed.
Flow Field in a Single-Stage Model Air Turbine Rotor-Stator Cavity With Pre-Swirled Purge Flow
Authors: Dunn D. M. , Zhou D. W. , Saha K. , Squires K. D. , Roy R. P. , Kim Y. W. , Moon H. K.
Published: January 2011

The present work applies a combined approach of experimental measurement and numerical simulation of the flow in a single-stage model gas turbine. Mainstream gas ingestion into the cavity is reduced by utilizing two axially overlapping seal rings, one on the rotor disk and the other on the stator wall. Additionally, pre-swirled purge air is injected into the rotor-stator cavity through the stator radially inboard of the two seal rings. Flow field predictions from the simulations are compared against experimental measurements of static pressure, velocity, and tracer gas concentration acquired in a nearly identical model configuration. The experiment chosen for simulation was performed for a main airflow Reynolds number of 7.86×104 at a rotor disk speed of 3000rpm with a rotational Reynolds number of 8.74×105, and purge air nondimensional flow rate cw = 4806. The simulation models a 1/14 sector of the cylindrical geometry consisting of four rotor blades and four stator vanes. Gambit 2.4.6 was used to generate the three-dimensional grids ranging from 10 to 20 million elements. Effects of turbulence were modeled using the single-equation Spalart-Allmaras as well as the realizable k-epsilon models. Computations were performed using FLUENT 6.3 for both a simplified frozen-rotor formulation (steady) and a subsequent time-dependent (unsteady) computation. The simulation results show symmetry in the pressure and velocity distributions across each pitch of the stator vanes, but larger scale structures across all four blade pitches are also realized. Velocity distributions were scrutinized in the rotor-stator cavity and are in reasonable agreement with the measurements. Static pressure and tangential velocity components are accurately predicted.
Numerical Predictions of the Effect of Rotation on Fluid Flow and Heat Transfer in an Engine-Similar Two-Pass Internal Cooling Channel With Smooth and Ribbed Walls
Authors: Schu¨ler M. , Dreher H.-M. , Neumann S. O. , Weigand B. , Elfert M.
Published: January 2011

In the present study, a two-pass internal cooling channel with engine-similar cross-sections was investigated numerically. The channel featured a trapezoidal inlet pass, a sharp 180° bend and a nearly rectangular outlet pass. Calculations were conducted for a configuration with smooth walls and walls equipped with 45° skewed ribs (P/e = 10, e/dh = 0.1) at a Reynolds number of Re = 50,000. The present study focused on the effect of rotation on fluid flow and heat transfer. The investigated rotation numbers were Ro = 0.0 and 0.10. The computations were performed by solving the Reynolds-averaged Navier-Stokes equations (RANS method) with the commercial Finite-Volume solver FLUENT using a low-Re k-ω-SST turbulence model. The numerical grids were block-structured hexahedral meshes generated with POINTWISE. Flow field measurements were independently performed at DLR using Particle Image Velocimetry. In the smooth channel rotation had a large impact on secondary flows. Especially, rotation induced vortices completely changed the flow field. Rotation also changed flow impingement on tip and outlet pass side wall. Heat transfer in the outlet pass was strongly altered by rotation. In contrast to the smooth channel, rotation showed less influence on heat transfer in the ribbed channel. This is due to a strong secondary flow field induced by the ribs. However, in the outlet pass Coriolis force markedly affected the rib induced secondary flow field. The influence of rotation on heat transfer was visible in particular in the bend region and in the second pass directly downstream of the bend.
Parametrical Investigation of Turbine Stages With Open Tip Clearance for the Purpose of Stage Efficiency Increase
Authors: Granovskiy Andrei , Kostege Mikhail , Lomakin Nikolay
Published: January 2011

The aerodynamic loss due to tip leakage vortex of rotor blades represents a significant part of viscous losses in axial flow turbines. The mixing of leakage flow with the main rotor passage flow causes losses and reduces turbine stage efficiency. Many measures have been proposed to reduce the loss in the tip clearance area. In this paper the reduction of the tip clearance loss due to changes made to the blade tip section profile is presented. The blade tip profile was modified to decrease the pressure gradient between pressure surface and suction surface. This approach allows the reduction of tip leakage and tip vortex strength and consequently the reduction of tip clearance losses. A 3D Navier-Stokes solver with q-ω turbulence model is used to analyze the flow in the turbine with various tip section profiles. Test data of three single-stage experimental turbines have been used to validate analytical predictions: • Highly loaded turbine stage with a pressure ratio π0T = 3.2 and reaction degree ρmean = 0.5. • Two turbines with a pressure ratio π0T = 3.9. (One with high degree of reaction ρmean = 0.55; the other with low degree of reaction ρmean = 0.26). The numerical investigation of the influence of various tip section profiles on stage efficiency was carried out in the range of relative tip clearance 0.5%–2.4% with the objective of a decreasing the influence of the tip clearance on the stage efficiency.
Numerical Re-Design of a Heavy Duty Gas Turbine Hydrogen-Fired Combustion Chamber
Authors: Marini Alessandro , Riccio Giovanni , Martelli Francesco , Sigali Stefano , Cocchi Stefano
Published: January 2011

The present work describes the numerical methodology followed to characterize and study modifications to a silo-type diffusive combustion chamber installed on a GE10, a 10 MW class heavy-duty gas turbine manufactured by GE Oil & Gas. The goal of the work was to investigate modifications to the combustion chamber to allow operation with 100% hydrogen fuel at reduced NOx production in dry conditions. The investigation focused mainly on the burner; the liner was not substantially changed. The swirler and the fuel injection holes were redesigned to achieve better fuel-air mixing and a higher airflow rate in the primary zone of the combustor, maintaining a diffusion flame scheme. The proposed modifications were analyzed using a 3D CFD RANS reactive procedure based on commercial codes. The method was previously validated by comparison with the experimental data from the full scale tests performed at the Enel Facility at Sesta, Italy. In-house codes were developed for the post-processing of the results. The numerical analysis has shown that the modified version can provide a NOx reduction up to 40%. The results are discussed focusing on the effect of fuel injection scheme on mixing quality and NOx emission containment.
The Application of Low-Profile Vortex Generators in an Intermediate Turbine Diffuser
Authors: Santner C. , Go¨ttlich E. , Marn A. , Hubinka J. , Paradiso B.
Published: January 2011

The demand of further increased bypass ratios for turbofan engines will lead to low pressure turbines with larger diameter and lower rotational speed in conventional high-bypass aero engine architectures. Due to that, it is necessary to guide the flow leaving the high pressure turbine to the low pressure turbine at a larger diameter without any severe loss generating separation or flow disturbances. To reduce costs and weight this turbine duct has to be as short as possible. This results in super-aggressive (very high diffusion) s-shaped geometries where the boundary layers are not able to withstand the strong adverse pressure gradient which results in flow separation. This paper describes the flow through a fully separated duct as a baseline configuration. In a next step the influence of passive flow control devices onto this separation has been investigated. Therefore, low-profile vortex generators were applied within the first bend of this s-shaped intermediate turbine diffuser in order to energize the boundary layer and further reduce or even suppress the occurring separation. This configuration was investigated downstream a transonic turbine stage. Measurements were performed by means of five-hole-probes, static wall pressure taps and oil flow visualization at the duct endwalls. For a better understanding of the flow behavior the vortex generators were also investigated in a two-dimensional rectangular s-shaped duct using the same Mach number level. Results showed that the vortex generators reduce the separation in the 2D-duct but have no distinct influence on the separation within the turbine duct due to wakes as well as strong secondary flow effects. This work is part of the European project AIDA (Contract: AST3-CT-2003-502836).
The Flow Field Within an Axial Flow Compressor at Extremely High Flow Coefficients
Authors: Gill A. , von Backstro¨m T. W. , Harms T. M.
Published: January 2011

This article describes an experimental investigation of the flow structures occurring in a three-stage axial flow compressor during fourth quadrant operation in the incompressible flow regime. In fourth quadrant operation, the flow coefficient exceeds the design value to such a degree that the pressure difference between the compressor inlet and outlet becomes negative, and the compressor acts as a badly designed turbine. The pressure rise characteristic curve thus extends into the fourth quadrant of the compressor map. A three stage axial flow compressor, with a mass flow rate of 2.7 kg/s and a pressure ratio of 1.022 was investigated. The design rotor tip Mach number is 0.2. Three operational points within the fourth quadrant were investigated, at flow coefficients of 0.665, 0.747 and 1.024. A five hole conical probe and a 50 μm diameter inclined hot film anemometer were used in this investigation. Radial traverses downstream of rotor rows and a time-dependent area traverse downstream of the first stage stator were performed. Three-dimensional steady-state and time-dependent numerical Navier-Stokes solutions for single blade passages in each blade row for each of the cases are compared with experimental work. Large wakes were observed downstream of all stator rows, as a result of significant flow separation on stator blades. The area fraction of the flow passage affected by the wakes increases as the flow coefficient increases. Flow through rotor blade-passages is heavily affected by the blade position relative to upstream stator wakes. Due to the effect of the stator wakes on downstream blading, time-dependent solutions using the nonlinear harmonic approximation were found to agree better with experimental results than steady-state solutions using mixing planes between blade rows.
Predictive Emissions Monitoring Using a Continuously Updating Neural Network
Authors: Vanderhaegen Evert , Deneve Michae¨l , Laget Hannes , Faniel Nathalie , Mertens Jan
Published: January 2011

In the European Union, power plants of more than 50 MW (thermal energy) need to comply with the Large Combustion Plant Directive (LCPD, 2001) implying that flue gas emissions need to be measured continuously. Traditionally, emissions from power plants are measured using Automated Measuring Systems (AMS). The LCPD states that no more than 10 days of emission data may be lost within one year including days needed for maintenance. This is the reason why more and more power plants are currently installing a second, back-up AMS since they have problems with the availability of their AMS. Since early 1990’s, Predictive Emissions Monitoring Systems (PEMS) are being developed and accepted by some local authorities within Europe and the United States. PEMS are in contrast to AMS based on the prediction of gaseous emissions (most commonly NOx and CO) using plant operational data (eg. fuel properties, pressure, temperature, excess air, …) rather than the actual measurement of these emissions. The goal of this study is to develop a robust PEMS that can accurately predict the NOx and CO emissions across the entire normal working range of a gas turbine. Furthermore, the PEMS should require as little maintenance as possible. The study does not intend to replace the AMS by a PEMS but rather to use the PEMS as a backup for the AMS. Operational data of a gas turbine, acquired over a long period, was used to identify inputs with a high influence on the NOx and CO formation. Consequently, simulations were done testing different model structures and calibration methodologies. The study shows that a static model failed to predict the emissions accurately over long time periods. In contrast, a dynamic or self-adapting algorithm proved to be most efficient in predicting the emissions over a long time period with a minimum of required intervention and maintenance. The self-adapting algorithm uses measured AMS data to continuously update the neural network. Since the PEMS is developed as a backup for the AMS, these data are readily available. The study shows that in case of a failing AMS, the developed model could accurately predict the NOx emissions for a duration of several weeks. Although not discussed in detail in this study, a quality assurance system of the PEMS is also developed since the PEMS needs to comply to the EN14181 (as does any AMS). The PEMS as a backup of the AMS instead of a second AMS is cost and time saving. Not only is the purchase of a second AMS avoided (between 40 and 100 k€) but equally important and of the same order of magnitude are the cost and time savings with respect to the Quality Assurance of the second AMS.
Transient Numerical Investigation of Rotor Clocking in 1.5 Stage of an Axial Test Turbine With a Blade-to-Vane Ratio of 1.5
Authors: Kusterer K. , Moritz N. , Bohn D. , Sugimoto T. , Tanaka R.
Published: January 2011

In recent years clocking of stators and rotors in multi-stage axial compressors and turbines has become an important scope of scientific investigations aiming to reduce aerodynamic losses in these turbo machines and to affect the aerodynamic and thermal load, the unsteady forces and the flow conditions of following stages. Regarding the loss reduction, the target of clocking is to find a relative circumferential position of successive rows of vanes or blades, so that the low momentum region in the wake of the upstream airfoil impinges the leading edge of the vanes or blades of the following stage. Numerical investigations of clocking in a multi-stage arrangement require the transient calculation of the instationary flow field, so that the transient interaction between stationary and rotating airfoils is part of the calculations. In the present investigation, the transient version of the conjugate fluid flow and heat transfer solver CHTflow has been used. In this theoretical study the 1st rotor, 2nd vane and 2nd rotor of an axial test turbine are investigated. The cooling of 1st rotor and 2nd vane is taken into account by slot injections on the blade surface. The influence of the blade clocking has been investigated at four different circumferential positions of the 2nd rotor with respect to the 1st rotor. Despite of this change, all other geometric parameters and the boundary conditions are unchanged for the four configurations. The transient flow structures are analyzed by visualization of the entropy distribution and the helicity at different time steps in a midspan section. Thereby, it can be seen how the wakes of the rotor 1 are chopped by the successive stator and how they impinge on rotor 2. It is shown that the flow pattern in front of rotor 2 is not equivalent for the 3 blades in the calculated segment. Thus, the wakes of rotor 1 cannot continuously impinge the leading edges of rotor 2, so that there is no significant influence of the clocking position on the time averaged total efficiency of the stages to be found, although the time dependent values of the total efficiency show different behaviors for the analyzed clocking positions. Therefore it can be concluded that for the investigated configuration the effect of clocking on the efficiency is significantly reduced in case of a blade-to-vane ratio of 1.5, so that the distinction between advantageous and disadvantageous assembly positions is not important.
The Battle of CO2 Capture Technologies
Authors: Narula Ram G. , Wen Harvey
Published: January 2011

Coal is an abundant, widespread, cheap energy source and contributes to 39% of the world’s electric power generation. Coal releases large amounts of carbon dioxide (CO2), which is believed to play a major role in global warming and climate change. To de-carbonize power generation, three distinct carbon capture technologies are in varying stages of development. These include pre-combustion carbon capture through the use of integrated coal gasification combined cycle (IGCC), post-combustion carbon capture from a pulverized-coal (PC)-fired power plant flue gas using monoethanolamine (MEA) or ammonia (NH3), and oxy-combustion technology. In the latter technology, oxygen is first separated from nitrogen in an air separator unit and used for combustion of coal in a conventional PC boiler. With oxy-combustion technology, the resulting flue gas is predominantly CO2, which makes CO2 capture easier than in the PC-MEA case. This paper discusses the development status as well as the advantages, limitations, performance and economics of each technology in regard to the capture and non-capture cases.
Effect of Geometry Deviations on the Aerodynamic Performance of an Outlet Guide Vane Cascade
Authors: Chernoray Valery , Ore Sofia , Larsson Jonas
Published: January 2011

Influence of the surface geometry variations on the flow in a low-pressure turbine outlet guide vane cascade is studied experimentally and numerically. Experiments are performed in a linear cascade facility at Chalmers, and numerical simulations are carried out at the rig conditions using steady RANS equations. Investigated surface nonconformances are specially designed two- and three-dimensional surface roughness elements which simulate a generic welding trace and a surface repair patch. These surface modifications were implemented on the cascade vanes at different surface locations, and the aerodynamics of the cascade with geometry deviations was compared to the baseline case without them. Investigated characteristics include the cascade performance in terms of the total pressure loss and flow turning angle as well as a detailed description of the downstream development of the secondary flow field. It is found that there is a range of locations on the vane surface where even relatively large roughness elements as were investigated (with height up to 20% of the blade maximum thickness) do not affect the OGV performance significantly. Another range of locations, where the effect is critical and the flow separation is triggered, are identified as well. An accurate prediction of the separation margins in the latter case was found a challenging task for turbulence models.
An Efficient Procedure for the Analysis of Heavy Duty Gas Turbine Secondary Flows in Different Operating Conditions
Authors: Cerutti Matteo , Bozzi Luca , Bonzani Federico , Carcasci Carlo
Published: January 2011

Combined cycle and partial load operating of modern heavy-duty gas turbines require highly efficient secondary air systems to supply both cooling and sealing air. Accurate performance predictions are then a fundamental demand over a wide range of operability. The paper describes the development of an efficient procedure for the investigation of gas turbine secondary flows, based on an in-house made fluid network solver, written in Matlab® environment. Fast network generation and debugging are achieved thanks to Simulink® graphical interface and modular structure, allowing predictions of the whole secondary air system. A crucial aspect of such an analysis is the calculation of blade and vane cooling flows, taking into account the interaction between inner and outer extraction lines. The problem is closed thanks to ad-hoc calculated transfer functions: cooling system performances and flow functions are solved in a pre-processing phase and results correlated to influencing parameters using Response Surface Methodology (RSM) and Design of Experiments (DOE) techniques. The procedure has been proved on the secondary air system of the AE94.3A2 Ansaldo Energia gas turbine. Flow functions for the cooling system of the first stage blade, calculated by RSM and DOE techniques, are presented. Flow functions based calculation of film cooling, tip cooling and trailing edge cooling air flows is described in details.
Numerical Study of Aerodynamic Losses of Effusion Cooling Holes in Aero-Engine Combustor Liners
Authors: Andreini A. , Bonini A. , Caciolli G. , Facchini B. , Taddei S.
Published: January 2011

Due to the stringent cooling requirements of novel aeroengines combustor liners, a comprehensive understanding of the phenomena concerning the interaction of hot gases with typical coolant jets plays a major role in the design of efficient cooling systems. In this work an aerodynamic analysis of the effusion cooling system of an aero-engine combustor liner was performed; the aim was the definition of a correlation for the discharge coefficient (CD) of the single effusion hole. The data was taken from a set of CFD RANS simulations, in which the behavior of the effusion cooling system was investigated over a wide range of thermo fluid-dynamics conditions. In some of these tests, the influence on the effusion flow of an additional air bleeding port was taken in account, making possible to analyze its effects on effusion holes CD. An in depth analysis of the numerical data set has pointed out the opportunity of an efficient reduction through the ratio of the annulus and the hole Reynolds numbers: the dependence of the discharge coefficients from this parameter is roughly linear. The correlation was included in an in-house one dimensional thermo-fluid network solver and its results were compared with CFD data. An overall good agreement of pressure and mass flow rates distributions was observed. The main source of inaccuracy was observed in the case of relevant air bleed mass flow rates, due to the inherent three-dimensional behavior of the flow close to bleed opening. An additional comparison with experimental data was performed in order to improve the confidence in the accuracy of the correlation: within the validity range of pressure ratio in which the correlation is defined (> 1.02), this comparison pointed out a good reliability in the prediction of discharge coefficients. An approach to model air bleeding was then proposed, with the assessment of its impact on liner wall temperature prediction.
Development and Validation of a Model for Axial Compressor Fouling Simulation
Authors: Melino F. , Peretto A. , Spina P. R.
Published: January 2011

Gas turbine axial compressor performance are heavily influenced by blade fouling; as a result, the gas turbines efficiency and producible power output decrease. In this study a model, able to evaluate the performance degradation of an axial compressor due to fouling, is developed and validated. The model is validated against experimental results available in literature and included into a computer code developed by the Authors (IN.FO.G.T.E) which is able to estimate the performance of every commercial gas turbine by using a stage stacking methods for the simulation of compressor behavior. The goal of this study is to show and discuss the change in gas turbine main performance (such as efficiency, power output, compressor inlet mass flow rate, pressure ratio) due to compressor fouling and also highlight and discuss the change in compressor stages performance curves.
Large Tip Clearance Flows in Two Compressor Cascades
Authors: Williams Richard , Ingram Grant , Gregory-Smith David
Published: January 2011

Large tip clearances in the region of six percent span exist in the high pressure stages of compressors of industrial gas turbines. The over tip clearance flow causes significant blockage and accounts for the largest proportion of loss in the high pressure compressor. This paper examines large tip clearances in two different compressor cascades. The first was a cascade of controlled-diffusion blades and the second cascade had geometry more representative of modern engine practise. The second cascade also featured realistic inlet boundary layer conditions delivered by an upstream injection system. Increasing the understanding of such flows will allow for improvements in the design of such compressors. The key conclusions of this paper are: a) At large tip clearances the tip clearance is the primary variable influencing the flow pattern, blade geometry is a secondary consideration; b) The blade geometry has a significant influence on loss generation through the cascade; c) With increasing tip clearance the loss generation through the cascade can diminish but the flow turning is much reduced; and d) The blade loading at the tip can increase at large tip clearances.