bluuebunny/crossref_metadata_2025_split huggingface dataset:
Investigation of Different Surge Handling Strategies and its Impact on the Cogeneration Performance for a Single-Shaft Gas Turbine Operating on Syngas
Authors: Montemayor Javier Beltran , Axelsson Lars-Uno 
Published: August 2015

The increasing demand for decentralized power has led to a growing interest in smaller gas turbines for cogeneration applications. One benefit of decentralized power generation is the possibility to utilize fuels that are locally available. One example is syngas, which has gained increasing interest during the recent years. Compared to natural gas the syngas fuels have a large amount of dilutants, such as nitrogen and carbon dioxide, which results in a very low energy density. Hence, significant larger fuel flow is required. However, the added fuel flow will decrease the compressor surge margin and eventually drive the compressor towards surge. Several methods exist to address this issue including variable inlet guide vanes, increased turbine throat area, compressor bleeding and decreased combustor outlet temperature.
This paper examines the operability of a generic all-radial single-shaft gas turbine in the 2 MW power range when running on syngas with different heating values. The above methods to combat the decreased surge margin will be analyzed using detailed cycle simulations and their advantages and disadvantages will be discussed. It was found that the increased throat area is the most beneficial of the four methods. The net power output is nearly the same as for natural gas operation and the heat rate is the lowest of the four methods.
Analysis of Long-Term Gas Turbine Operation With a Model-Based Data Reconciliation Technique
Authors: Rudolf Christian , Wirsum Manfred , Gassner Martin , Bernero Stefano 
Published: August 2015

The continuous monitoring of gas turbines in commercial power plant operation provides long-term engine data of field units. Evaluation of the engine performance is challenging as, apart from variations of operating points and environmental conditions, the state of the engine is subject to changes due to the ageing of engine components. The measurement devices applied to the unit influence the analysis by means of their accuracy, which may itself alter with time. Furthermore, the available measurements do usually not cover all necessary information for the evaluation of the engine performance. To overcome these issues, this paper describes a method to systematically evaluate long term operation data without the incorporation of engine design models since the latter do not cover performance changes when components are ageing. Key focus of the methodology thereby is to assess long-term emission performance in the most reliable manner.
The analysis applies a data reconciliation method to long-term operating data in order to model the engine performance including non-measured variables and to account for measurement inaccuracies. This procedure relies on redundancies in the data set due to available measurements and the identification of suitable additional constituting equations that are independent of component ageing. The resulting over-determined set of equations allows for performing a data set optimization with respect to a minimal cumulated deviation to the measurement values, which represents the most probable, real state of the engine. The paper illustrates the development and application of the method to analyse the gas path of a commercial gas turbine in a combined cycle power plant with long-term operating data.
Aerodynamic Optimization of the Gas Outlet Casing of the Axial Turbine Stage in a Turbocharger With Experimental Validation
Authors: So K. K. , Sofia A. , Fischer M. , Phillipsen B. A. 
Published: August 2015

This paper presents a design process for optimizing the aerodynamic performance of the gas outlet casing on single-stage turbines with high specific flow capacity.
A full-annulus flange-to-flange (inlet-to-outlet) steady-state CFD model of the turbine stage which takes into account the interaction between the rotor and the gas outlet casing and ensures an accurate inflow condition for the latter, is used to predict the turbine stage performance. A meta-model based optimization for the gas outlet casing is then performed and a simplified CFD model is used for sampling and optimization. The geometry of the gas outlet casing is fully parameterized to enable the simultaneous variation of diffuser and collector geometry and an ordinal regression optimization algorithm is adopted for the objectives of maximizing the static pressure recovery of the gas outlet casing and ensuring the design robustness.
Extensive test measurement of the turbine stage with its baseline and optimized gas outlet casing geometries on a full scale turbocharger test bench validates the CFD results and confirms the significant improvement of the exhaust casing pressure recovery, which leads to an improvement of turbine efficiency between 1.3 and 2.4 % points over the relevant considered operating range. Traverse measurement using five-hole probes and the flow field predicted by CFD are in good agreement. Evaluation of the CFD results highlights a significant loss in the collector despite a high pressure recovery at the end of the diffuser for the baseline gas outlet casing. For the optimized geometry, the more uniform flow at the diffuser outlet results in greatly reduced loss in the 90° turn in the casing, and thus higher pressure recovery and turbine efficiency.
Heat Transfer and Pressure Drop Measurements on Rotating Matrix Cooling Geometries for Airfoil Trailing Edges
Authors: Carcasci Carlo , Facchini Bruno , Pievaroli Marco , Tarchi Lorenzo , Ceccherini Alberto , Innocenti Luca 
Published: August 2015

In the present paper the combined effects of rotation and channel orientation on heat transfer and pressure drop along two scaled up matrix geometries suitable for trailing edge cooling of gas turbine airfoils are investigated.
Experimental tests were carried out under static and rotating conditions. Rotating tests were performed for two different orientations of the matrix channel with respect to the rotating plane: 0deg and 30deg. This latter configuration is representative of the exit angle of a real gas turbine blade. Test models are designed in order to replicate an internal geometry suitable for blade trailing edge cooling, with a 90deg turning flow before entering the matrix array which has an axial development.
Both the investigated geometries have a cross angle of 45deg between ribs and different values of sub-channels and rib thickness: one has four sub-channels and lower rib thickness (open area 84.5%), one has six sub-channels and higher rib thickness (open area 53.5%). Both geometries have a converging angle of 11.4deg.
Matrix models have been axially divided in 5 aluminum elements per side in order to evaluate the heat transfer coefficient in 5 different locations in the main flow direction. Metal temperature was measured with embedded thermocouples and thin-foil heaters were used to provide a constant heat flux during each test.
Heat transfer coefficients were measured applying a steady state technique based on a regional average method and varying the sub-channel Reynolds number Res from 2000 to 10000 and the sub-channel Rotation number Ros from 0 to 0.250 in order to have both Reynolds and Rotation number similitude with the real conditions.
A post-processing procedure, which takes into account the temperature gradients within the model, was developed to correctly compute average heat transfer coefficients starting from discrete temperature measurements.
Lumped Capacitance and 3D CFD Conjugate Heat Transfer Modelling of an Automotive Turbocharger
Authors: Burke R. , Copeland C. , Duda T. , Reyes-Belmonte M. A. 
Published: August 2015

One dimensional wave-action engine models have become an essential tool within engine development including stages of component selection, understanding system interactions and control strategy development. Simple turbocharger models are seen as a weak link in the accuracy of these simulation tools and advanced models have been proposed to account for phenomena including heat transfer. In order to run within a full engine code, these models are necessarily simple in structure yet are required to describe a highly complex 3D problem. This paper aims to assess the validity of one of the key assumptions in simple heat transfer models, namely, that the heat transfer between the compressor casing and intake air occurs only after the compression process.
Initially a sensitivity study was conducted on a simple lumped capacity thermal model of a turbocharger. A new partition parameter was introduced αA, which divides the internal wetted area of the compressor housing into pre and post compression. The sensitivity of heat fluxes to αA was quantified with respect to the sensitivity to turbine inlet temperature (TIT). At low speeds, the TIT was the dominant effect on compressor efficiency whereas at high speed αA had a similar influence to TIT. However, modelling of the conduction within the compressor housing using an additional thermal resistance caused changes in heat flows of less than 10%.
Three dimensional CFD analysis was undertaken using a number of cases approximating different values of αA. It was seen that when considering a case similar to αA=0, meaning that heat transfer on the compressor side is considered to occur only after the compression process, significant temperature could build up in the impeller area of the compressor housing, indicating the importance of the pre-compression heat path. The 3D simulation was used to estimate a realistic value for αA which was suggested to be between 0.15 and 0.3. Using a value of this magnitude in the lumped capacitance model showed that at low speed there would be less than 1% point effect on apparent efficiency which would be negligible compared to the 8% point seen as a result of TIT. In contrast, at high speeds, the impact of αA was similar to that of TIT, both leading to approximately 1% point apparent efficiency error.
The Influence of Tip Clearance on Supercritical CO2 Centrifugal Compressor Performance
Authors: Zhao Hang , Deng Qinghua , Zhang Hanzhen , Feng Zhenping 
Published: August 2015

The supercritical CO2 (SCO2) compressor is the key component of the SCO2 Brayton cycle. It is considered as one of the most promising power conversion systems, because the compressor could consume small compression work as operating condition near the critical point. In this paper, a researched SCO2 compressor was designed to operate with slightly above the vapor-liquid critical point of CO2. The flow characteristics were investigated by NUMECA FINE/Turbo, coupled with the thermophysical properties of CO2 in REFPROP database. Particular attention was paid on the blade tip clearance flow characteristics. The formation and development mechanism of the blade tip low pressure regions and the parameters distributional difference were studied. Then, the effects of different blade tip clearance types on the low pressure regions of the blade leading edge tip and the aerodynamic performance of SCO2 compressor stage were compared and discussed.
The combined effects of the injecting action, the blade leading edge tip shedding vortex, and the flow acceleration can lead the fluid thermodynamic state to enter the low pressure and low temperature regions at the leading edge of both the main blade and splitter blade tip. Especially, the condensation maybe exist at sharp corner of the blade tip leading edge. However, the tip clearance leakage vortex can significantly inhibit and weaken the low pressure regions. With the increase of tip clearance, the distribution and development tendency of the static pressure and temperature on the main blade leading edge tip are similar, but there is a significant influence for the splitter blades tip. Furthermore, because the tip clearance increases, the SCO2 compressor stage pressure ratio and isentropic efficiency have degraded in whole operating range, but the stable operation range is expanded. Therefore, the above comprehensive factors should be considered during the design process in order to select the suitable tip clearance shape.
A Quasi-First Order Optimization Method and its Demonstration on the Optimization of a U-Bend
Authors: Akin M. Bugra , Sanz Wolfgang 
Published: August 2015

Optimal shape design is widely used today to improve a variety of designs. It is a challenging task and several methods have been developed. These methods are generally classified by the order of derivatives used. They are zero, first and second order methods, which, as their names imply, use only the function values, first and second order derivatives, respectively. There are two common approaches to first order methods. These are the finite difference method and the adjoint method. The finite difference method requires an additional CFD calculation for each parameter, which quickly becomes computationally very expensive as the number of parameters rise. The adjoint method provides a computationally efficient alternative in such cases. But the computational cost of the adjoint method also becomes expensive if additional constraints are introduced or when multi-objective optimizations are considered.
This paper presents a novel optimization strategy which can be classified as a quasi-gradient based optimization method. As with the finite differences method an additional CFD calculation is performed for each parameter. But in order to save computational time the simulations are not performed to full convergence so that the derivatives are not calculated accurately. The only information that can be obtained in this way is whether the chosen contour manipulation leads to an improvement. A line search method is introduced that can find an optimum using this incomplete gradient information. The optimization method is demonstrated by the quasi-3d optimization of a U-bend.
Automated Calibration of Compressor Loss and Deviation Correlations
Authors: Schnoes Markus , Nicke Eberhard 
Published: August 2015

Throughout preliminary design of multistage axial compressors performance prediction is based on fast 2D throughflow codes incorporating models for loss, deviation and secondary flow effects. To achieve high solution quality these models have to be calibrated on experimental or simulation data. Particularly 2D cascade loss and deviation correlations are of major importance. Correlations available in literature are mostly based on data from early cascade wind tunnel tests of NACA-65 and DCA profile families. In general, calibrations for state-of-the-art airfoil shapes are not published.
This paper describes a methodology to re-calibrate parameters of well-known empirical loss and deviation correlations to CFD simulations of compressor cascades. Correlations describing choke incidence, on- and off-design loss and deviation are presented, resolving dependencies on inlet flow angle, Mach number, Reynolds number and MVDR. Based on a set of loss and deviation polars, nonlinear least-squares problems are solved to determine the empirical parameters. By including shock losses and choking incidence, the model is suited for subsonic as well as supersonic inlet conditions.
Results of the calibration process are shown for two state-of-the-art compressor cascades. The overall model is evaluated on a sub- and a transonic compressor as well as on a transonic fan comparing throughflow results with 3D CFD and experimental data. A high level of automation makes the methodology applicable to arbitrary profile families or in combination with programs for parametric airfoil design.
Advanced Turbine Preliminary Design Environment for the Automatic Generation of Secondary Air System Models
Authors: Villazon Jose Maria Rey , Wildow Toni , Benton Robert , Eydam Tobias , Kühhorn Arnold 
Published: August 2015

The design and development process of an aero engine is a complex and time-consuming task that involves many disciplines and company departments with different objectives and requirements. Along the preliminary design phase, multiple concepts are assessed in order to select a competitive technology. The engine design process, which was traditionally subdivided into modular component tasks, is nowadays considered as a multi-disciplinary workflow. Having recognized the need for developing advanced turbine preliminary design tools, this work focuses on enhancing the integration of turbine design disciplines, improving the accuracy of models and speeding the time to generate models.
The proposed process facilitates an automated turbine Secondary Air System (SAS) and turbine discs concept definition. Furthermore, the process of CAD models and flow network models generation is accelerated via automation of the engineering workflow. This is accomplished through a novel Java based data model, where the design of turbine discs and SAS features is captured in a programmable framework.
In the application section, the preliminary design definition of a reference HP turbine subsystem is replicated using the newly developed common design environment. The automated workflow is then used to generate the corresponding CAD models, recognize the subsystem flow network, and generate the 1D flow network model. The results are then compared to the experimentally validated model of a reference engine. As conclusion, the automated workflow offers a quick and parametric model generation process, while providing a good level of fidelity for the preliminary design phase.
Frequency Domain Predictions of Acoustic Wave Propagation and Losses in a Swirl Burner With Linearized Navier-Stokes Equations
Authors: Zahn Max , Schulze Moritz , Hirsch Christoph , Betz Michael , Sattelmayer Thomas 
Published: August 2015

A low-cost, high-quality hybrid CFD/CAA-methodology is used to predict the acoustic properties of a swirl burner including its complex swirl flow conditions. The numerically determined burner transfer matrix is validated against experimental data. The results demonstrate the capability of this low-cost hybrid approach to predict the acoustic characteristics of combustor components with high geometrical complexity. Most importantly it captures the effect of mean flow quantities on the fluctuating field. This causes the loss of acoustic energy and thus constitutes sources of acoustic damping. In this regard, reliable data can be obtained to characterize complex acoustic components at relatively low computational cost. Therefore, experimental efforts can be reduced which are generally required to provide data e.g. to set up low-order network models.
The insight into the field of fluctuating quantities allows the analysis of linear acoustic damping phenomena. Essentially, in the context of isentropic conditions acoustic energy is lost due to the formation of vorticity disturbances. Source regions for vorticity disturbances are identified at flow separation edges and within the multiple shear layers of the complex swirl flow.
Experimental Results of the First Two Stages of an Advanced Transonic Core Compressor Under Isolated and Multi-Stage Conditions
Authors: Prahst Patricia S. , Kulkarni Sameer , Sohn Ki H. 
Published: August 2015

NASA’s Environmentally Responsible Aviation (ERA) Program calls for investigation of the technology barriers associated with improved fuel efficiency for large gas turbine engines. Under ERA, the highly loaded core compressor technology program attempts to realize the fuel burn reduction goal by increasing overall pressure ratio of the compressor to increase thermal efficiency of the engine. Study engines with overall pressure ratio of 60 to 70 are now being investigated. This means that the high pressure compressor would have to almost double in pressure ratio while keeping a high level of efficiency.
NASA and GE teamed to address this challenge by testing the first two stages of an advanced GE compressor designed to meet the requirements of a very high pressure ratio core compressor. Previous test experience of a compressor which included these front two stages indicated a performance deficit relative to design intent. Therefore, the current rig was designed to run in 1-stage and 2-stage configurations in two separate tests to assess whether the bow shock of the second rotor interacting with the upstream stage contributed to the unpredicted performance deficit, or if the culprit was due to interaction of rotor 1 and stator 1. Thus, the goal was to fully understand the stage 1 performance under isolated and multi-stage conditions, and additionally to provide a detailed aerodynamic data set for CFD validation. Full use was made of steady and unsteady measurement methods to understand fluid dynamics loss source mechanisms due to rotor shock interaction and endwall losses.
This paper will present the description of the compressor test article and its measured performance and operability, for both the single stage and two stage configurations. We focus the paper on measurements at 97% corrected speed with design intent vane setting angles.
Design Improvement Survey for NOx Emissions Reduction of a Heavy-Duty Gas Turbine Partially Premixed Fuel Nozzle Operating With Natural Gas: Numerical Assessment
Authors: Innocenti Alessandro , Andreini Antonio , Facchini Bruno , Cerutti Matteo , Ceccherini Gianni , Riccio Giovanni 
Published: August 2015

A numerical investigation of a low NOx partially premixed fuel nozzle for heavy-duty gas turbine applications is presented in this paper. Availability of results from a recent test campaign on the same fuel nozzle architecture allowed the exhaustive comparison study presented in this work.
At first, an assessment of the turbulent combustion model was carried out, with a critical investigation of the expected turbulent combustion regimes in the system and taking into account the partially premixed nature of the flame due to the presence of diffusion type pilot flames. In particular, the Fluent partially premixed combustion model and a flamelet approach are used to simulate the flame. The laminar flamelet database is generated using the Flamelet Generated Manifold (FGM) chemistry reduction technique. Species and temperature are parameterized by mixture-fraction and progress variable. Comparisons with calculations with partially premixed model and the steady diffusion flamelet (SDF) database are made for the baseline configuration in order to discuss possible gains associated with the introduced dimension in the FGM database (reaction progress) which makes it possible to account for non-equilibrium effects. Numerical characterization of the baseline nozzle has been carried out in terms of NOx. Computed values for both the baseline and some alternative premixer designs have been then compared with experimental measurements on the reactive test rig at different operating conditions and different split ratios between main and pilot fuel. Numerical results allowed pointing out the fundamental NOx formation processes, both in terms of spatial distribution within the flame and in terms of different formation mechanisms. The obtained knowledge would allow further improvement of fuel nozzle design.