bluuebunny/crossref_metadata_2025_split huggingface dataset:
Large-Eddy Simulation and Linear Acoustic Analysis of a Diffusion Swirling Flame Under Forcing and Self-Excitation
Authors: Zhang Man 
Published: September 2016

A diffusion swirling flame under external forcing and self-excitation within a single swirler combustor have been studied in this paper with the large-eddy simulation and linear acoustic method. The combustor features pre-vaporized kerosene as the fuel, a single radial air swirler for flame stabilization and a square cross section chamber with adjustable length. Firstly, self-sustained pressure oscillation has been achieved by using of a chocked nozzle on the chamber outlet with large-eddy simulation. Dynamic pressure oscillations are analyzed in frequency domain through Fast Fourier Transform. The major pressure oscillation is identified as the 1st order longitudinal mode of the chamber. Further, the same frequency in the form of harmonic velocity oscillation is imposed on the inlet of the combustor while the chamber length has been changed. Based on this approach, a comparative study of the flame response with different excitation method but same frequency is carried out.
In both self-excited and forced cases, global and local flame responses as well as Rayleigh index have been analyzed and compared. With the flame response function, the excited acoustic modes under the influence of dynamic heat release have been predicted with linear acoustic method and compared with the results obtained from large-eddy simulation. Results show that the flame response presents a great difference in the spacial distribution with different excitation approaches. Thermo-acoustic interaction distributes along the flame front with the expansion of the flame under self-excitation while it damps with the acoustic propagating downstream under forcing condition. As the ratio of flame length to acoustic wave length could not be neglected for the diffusion swirling flame, the global flame response under forcing cannot represent the local response feature of the flame accurately, thus influencing the instability prediction.
NovaLT16 Bearing #2 Oil System Design: Test Rig Architecture, Functional Tests and Design Challenges
Authors: Mazzola Manfredi , Signori Andrea , Evangelisti Silvia , Di Sisto Paolo 
Published: September 2016

This paper describes the test rig and the functional tests that have been carried out to define a satisfying solution for the bearing #2 of NovaLT16, the GE Oil & Gas latest high-efficiency Gas Turbine.
The bearing #2 is a journal bearing located inside the compressor discharge casing and for this reason it is exposed to high temperatures. As a consequence, the lube oil provided to the bearing should not only lubricate the bearing pads but also cool the bearing housing and for this reason it should be more than what required by the pads. This large amount of oil is gathered by two bearing housing sumps and then drained by gravity through two pipes. Scavenge pumps are not required. The bearing housing also includes two circumferential labyrinth seals able to prevent oil leakage through the gap between the bearing housing and rotor. Both seals are purged by air extracted from the gas turbine compressor; the air not only creates a barrier to oil leakage but facilitate the draining inside the sumps.
An insufficient draining system may increase the oil levels in the bearing housing sumps above the safety limit and therefore it may generate oil leakage through the labyrinth seals. If the leakage reaches the external high-temperature zone then it can lead to oil degradation, decrease in GT performances and engine tripe due to smoke in the GT package in the worst case.
The draining system was therefore tested with an advanced test rig. The test rig was a 1:1 scale including the bearing housing, the bearing with its pads and also a mock-up of the shaft. The shaft was driven by an electric motor and therefore it was possible to verify the full operating speed range of the gas turbine. The gas turbine discharge compressor casing environment was simulated through electric resistances able to provide heat all around the bearing housing.
During the functional tests there have been simulated design and off-design conditions including extreme values of oil pressure, oil temperature, pitch and roll angles (for offshore application), shaft speed and buffering air. The results were mainly evaluated in terms of sump oil levels, shaft temperatures and bearing housing temperatures.
The full test campaign allowed to identify and validate a satisfying configuration both for on-shore and off-shore applications and to properly tune all the main bearing parameters in order to guarantee a robust configuration.
The NovaLT16 test campaign is part of the continue evolution in the GE O&G bearing #2 area design, started with the previous gas turbines (such as MS5002E and GE10).
The Design and Model Simulation of a Micro Gas Turbine Combustor Supplied With Methane/Syngas Fuels
Authors: Liu Chi-Rong , Shih Hsin-Yi 
Published: September 2016

The design and model simulation of a can combustor has been made for future syngas (mainly H2/CO mixtures) combustion application in a micro gas turbine. In previous modeling studies with methane as the fuel, the analysis indicated the design of the combustor is quite satisfactory for the 60-kW gas turbine; however, the cooling may be the primary concerns as several hot spots were found at the combustor exit. When the combustor is fueled with methane/syngas mixtures, the flames would be pushed to the sides of the combustor with the same fuel injection strategy. In order to sustain the power load, the exit temperature became too high for the turbine blades, which deteriorated the cooling issue of the compact combustor. Therefore, the designs of the fuel injection are modified, and film cooling is employed. Consequently, the simulation of the modified combustor is conducted by the commercial CFD software Fluent. The computational model consists of the three-dimensional, compressible k-ε model for turbulent flows and PPDF (Presumed Probability Density Function) model for combustion process between methane/syngas and air invoking a laminar flamelet assumption. The flamelet is generated by detailed chemical kinetics from GRI 3.0. Thermal and prompt NOx mechanisms are adopted to predict the NO formation.
At the designed operation conditions, the modeling results show that the high temperature flames are stabilized in the center of the primary zone where a recirculation zone is generated for methane combustion. The average exit temperature of the modified can combustor is 1293 K, which is close to the target temperature of 1200 K. Besides, the exit temperatures exhibit a more uniform distribution by coupling film cooling, resulting in a low pattern factor of 0.22. The NO emission is also low with the increased number of the dilution holes. Comparing to the results for the previous combustor, where the chemical equilibrium was assumed for the combustion process, the flame temperatures are predicted lower with laminar flamelet model. The combination of laminar flamelet and detailed chemistry produced more reasonable simulation results. When methane/syngas fuels are applied, the high temperature flames could also be stabilized in the core region of the primary zone by radially injecting the fuel inward instead of outward through the multiple fuel injectors. The cooling issues are also resolved through altering the air holes and the film cooling. The combustion characteristics were then investigated and discussed for future application of methane/syngas fuels in the micro gas turbine. Although further experimental testing is still needed to employ the syngas fuels for the micro gas turbine, the model simulation paves an important step to understand the combustion performance and the satisfactory design of the combustor.
Design Considerations for Tip Clearance Control and Measurement on a Turbine Rainbow Rotor With Multiple Blade Tip Geometries
Authors: Lavagnoli S. , De Maesschalck C. , Andreoli V. 
Published: September 2016

The accurate design, control and monitoring of the running gaps between static and moving components is vital to preserve the mechanical integrity and ensure the correct functioning of any compact rotating machinery. Throughout engine service, the rotor tip clearance undergoes large variations due to installation tolerances or as the result of different thermal expansion rates of the blades, rotor disk and casing during speed transients. Hence, active tip clearance control concepts and engine health monitoring systems rely on precise real-time gap measurements. Moreover, this tip gap information is crucial for engine development programs to verify the mechanical and aerothermal design, and validate numerical predictions.
This paper presents an overview of the critical design requirements for testing engine-representative blade tip flows in a rotating turbine facility. The manuscript specifically focuses on the challenges related with the design, verification and monitoring of the running tip clearance during a turbine experiment.
In the large-scale turbine facility of the von Karman Institute, a rainbow rotor was mounted for simultaneous aerothermal testing of multiple blade tip geometries. The tip shapes are a selection of high-performance squealer-like and contoured blade tip designs. On the rotor disc, the blades are arranged in seven sectors operating at different clearance levels from 0.5 up to 1.5% of the blade span.
Prior to manufacturing, the blade geometry was modified to compensate for the radial deformation of the rotating assembly under centrifugal loads. A numerical procedure was implemented to minimize the residual unbalance of the rotor in rainbow configuration, and to optimize the placement of every single airfoil within each sector. Subsequently, the rotor was balanced in-situ to reduce the vibrations and satisfy the international standards for high balance quality. The single-blade tip clearance in rotation was measured by three fast-response capacitive probes located at three distinct circumferential locations around the rotor annulus. Additionally, the minimum running blade clearance is captured with wear gauges located at five axial positions along the blades chord. The capacitance probes are self-calibrated using a multi-test strategy at several rotational speeds. The in-situ calibration methodology and dedicated data reduction techniques allow the accurate measurement of the distance between the turbine casing and the local blade tip features (rims and cavities) for each rotating airfoil separately. General guidelines are given for the design and calibration of a tip clearance measurement system that meets the required measurement accuracy and resolution in function of the sensor uncertainty, nominal tip clearance levels and tip seal geometry.
Effect of Aromatics in Jet Fuels on Spray Characteristics Downstream of an Aeronautical Pressure Swirl Atomizer
Authors: Leparoux Julien , Lecourt Renaud , Penanhoat Olivier 
Published: September 2016

Standard aeronautic fuels have a lower limit of aromatics of 8% (by volume) with about 18% for regular Jet A1. It has been shown that aromatics contained in Jet fuel have an impact on the fine particle emissions. In order to reduce these emissions, alternative fuels with lower aromatic content have been identified as a promising solution. Change Jet fuel composition can have several effects on spray and combustion behaviors, among others: atomization process, droplet evaporation, flame structure, pollutant and particle emissions. Then, it is necessary to evaluate the impact of this change on gas turbine performance and operability.
The present study is focused on the spray behavior investigation with different aromatic content. Four Jet fuels are investigated including conventional Jet A1 kerosene, drop-in fuel with a mixture of half conventional Jet fuel and synthetic paraffinic kerosene (SPK), SPK with 8% of aromatics and pure SPK. The tests are performed at atmospheric conditions on the MERCATO testbed located at ONERA (FR). Phase Doppler Anemometry (PDA) measurements are carried out for the four fuels on an injection system composed of a pressure swirl atomizer and an air swirler.
In this paper, a spray analysis of liquid velocity and droplet diameter measurements is described and linked to the variations of fuel properties. In the range of parameters covered by the four different fuels, it is shown that the spray behavior of each fuel is similar to the conventional Jet A1.
Numerical and Experimental Studies of Reynolds Number and Stator Clocking Effect on Performance of a High-Loaded Two-Stage Compressor With 3.7 Total Pressure Ratio
Authors: Mileshin V. , Druzhinin I. , Savin N. , Kozhemyako P. 
Published: September 2016

One of the vital tasks related to the improvement of the efficiency of turbomachines is an increase in their operating altitude by reducing the influence of Reynolds numbers, Re, on the turbomachine parameters. Therefore, the results presented in this work on the effect of Re on parameters of an axial compressor composed of two high-loaded stages are attractive both from scientific and practical points of view.
This work presents the results of experimental investigations of Re effect on gas-dynamic characteristics of a highly-loaded two-stage compressor (HPC-2), simulating the first two stages of the High Pressure Compressor, (HPC), for an advanced engine. The compressor has the following key gasdynamic parameters:
– corrected mass flow rate: 31.8 kg/s
– total pressure ratio: 3.7
– coefficient of aerodynamic load: 0.421
The experimental study of HPC-2 is carried out at the Central Institute of Aviation Motors, (CIAM), C-3 test facility. Reynolds number is changed by decreasing pressure at the inlet from 1 bar to 0.2 bar and changing the clocking position of stator rows at optimal points of performance for two rotational speeds, 0.7 and 0.88.
Tests of HPC-2 compressor show that a stepwise decrease of the inlet pressure from 1.0 0.7, 0.4 down to 0.2 bar (a decrease in Re from 1785000 to 294000) leads to a smooth decrease in maximum compressor efficiency by 1% and a shift of compressor characteristics towards lower air flow rates (by 2%).
The experimental studies of compressors for present-day engines show that unsteady processes play a key role in compressor efficiency and stability.
One way to control unsteady flow in compressors is clocking effect. Recently clocking of stators and rotors has been investigated in details using the HPC-2 two-stage compressor (total pressure ratio 3.7) in H = 0; M = 0 conditions. The effect of Reynolds number on compressor characteristics is studied in this work for the HPC-2 compressor while also investigating stator clocking effect. The rotor clocking effect is not studied. Tests of HPC-2 compressor show that a decrease in the inlet pressure and changes in Re results in a stronger clocking effect, which is 0.5% in terms of efficiency under atmospheric conditions at the inlet and reaches 1% with a decrease in the inlet pressure from atmospheric pressure down to 0.2 bar.
Effusion Cooled Combustor Liner Tiles With Modern Cooling Concepts: A Comparative Experimental Study
Authors: Jackowski Thomas , Schulz Achmed , Bauer Hans-Jörg , Gerendás Miklós , Behrendt Thomas 
Published: September 2016

Advanced combustion techniques in aero engines require highly effective cooling schemes of combustor liners. One parameter affecting the cooling performance is the geometry of the cooling holes themselves. So far, the freedom in the design of cooling holes was limited due to the manufacturing techniques. With emerging additive manufacturing methods, e.g. Direct Metal Laser Sintering, however, the geometry of the cooling holes is virtually unlimited. Especially the entrance and the curvature of the cooling holes determines the through-flow of the hole and consequently the cooling performance of the ejected cooling film.
In this study a set of combustor liner tiles with two innovative and four traditional cooling hole geometries will be analyzed and compared to each other in terms of cooling performance. The innovative geometries have bent cooling holes with a nearly horizontal outlet. All specimens have the same cooling hole pattern. The cooling performance is determined by comparing the total cooling effectiveness for a given pressure difference across the combustor liner tiles. The coolant mass flow rate is gained from experimentally determined discharge coefficients for the respective pressure difference. The first set of measurements is conducted in an atmospheric open-loop test rig at reduced temperatures but realistic density ratios between hot gas and coolant.
The specimen with the best cooling performance has been selected for an investigation in a high pressure test rig at realistic combustor conditions (pressure, temperature) including fluctuations of the cooling air to simulate combustion instabilities. The cooling performance again is determined by the total cooling effectiveness for a given pressure difference across the combustor liner tiles.
Analytical Uncertainty Analysis for Hot-Wire Measurements
Authors: Hösgen Christian , Behre Stephan , Hönen Herwart , Jeschke Peter 
Published: September 2016

This paper presents an analytical approach to quantify the uncertainty of velocity-, angle- and turbulence intensity measurements obtained by hot-wire anemometry. Based on the given results of the model, the paper points out the most critical parameters influencing the quality of the flow measurement.
The proposed approach is divided into three subsections, each accounting for the uncertainties due to the measurement equipment, probe calibration and measurements in the test-rig respectively. For every subsection the propagation of uncertainties is stated in the form of analytical formulas starting with the measurands, e.g. voltage, pressure, temperature, and ending with velocity, angle and turbulence intensity. The approach is applied to data, obtained by a triple hot-wire probe, on two-dimensional traverses in a 1.5-stage cold air turbine. The calibration process and, in particular, the temperature and pressure deviation between calibration and measurement are identified as having a major impact on the measurement uncertainty. With respect to the Aachen test-rig the uncertainty in velocity measurements is lower than 3.5%. The investigation of the turbulence intensity uncertainty yields an increase in uncertainty with increasing turbulence intensity but an insensitivity towards pitch- and yaw-angle uncertainty.
Numerical Simulation of the Flow in the Row Labyrinth Seal of an Axial Turbine Using a Low-Mach Preconditioning Technique
Authors: Fiedler Jens , Weber Anton , Bosco Arianna , Engel Karl , di Mare Francesca 
Published: September 2016

Labyrinth seals on shrouded blades are an effective way for reducing efficiency penalties, as compared to free ended blades. Due to the difficulties of gaining optical access to cavity regions, mostly pressure measurements are available in the literature, from which the details of the flow must be inferred. The use of numerical tools can provide insight in the flow topology and therefore help obtaining a better understanding of the factors (geometric, thermodynamic and aerodynamic) which can affect the performance of the machine. Whilst in the main passage relatively high Mach numbers are to be found (0.3–1.3), the flow field in the cavities is dominated by extremely low flow speeds with strong recirculation patterns. The treatments of such flows, where large disparities between the acoustic and convective speed exist, are known to be highly problematic if density-based solution methods are employed. As the flow conditions approach the incompressibility limit a degradation of the convergence behaviour can be observed, leading, potentially to incorrect solutions. In order to overcome these problems preconditioning methods can be conveniently applied to the Navier-Stokes equations. In the current work the formulation of a fully implicit local preconditioning method with domain control of the Mach number dependency is presented. Numerical simulations of turbomachine components are generally performed on truncated domains. In order to prevent unphysical reflections at open boundaries and interfaces non-reflecting boundary conditions have been developed, e.g. [6, 8]. As reported in the available literature, low Mach preconditioning can cause stability problems and strongly impair the quality of the results especially in proximity of the domain’s boundaries. As shown in [14, 21] an appropriate scaling treatment of the boundary conditions is also required to alleviate such issues.
In the current work non-reflecting boundary conditions, based on the formulation of Giles [6, 8], have been suitably modified to work reliably also in the limit of incompressible flows. To prove the robustness and accuracy of the algorithm implemented in the DLR’s CFD code TRACE, a canonical testcase representing an abstraction of the flow topology found in a labyrinth seal, such as a lid-driven cavity, is shown. Finally, the simulation of the steady flow in a multistage, shrouded low-pressure turbine is presented. For this, a classic RANS approach has been adopted using the k-ω model to illustrate the effectiveness of the developed method in a typical industrial application. Of particular significance and interest is the analysis of the mass conservation properties of the numerical scheme attained at mixing planes between rotor and stator and at non-matching grid interfaces, denoted as “zonal” and “zonal-mixed” interfaces.
Modulation and Radial Migration of Turbine Hub Cavity Modes by the Rim Seal Purge Flow
Authors: Schädler R. , Kalfas A. I. , Abhari R. S. , Schmid G. , Voelker S. 
Published: September 2016

In the present paper, the results of an experimental and numerical investigation of the hub cavity modes and their migration into the main annulus flow are presented. A one-and-a-half stage, unshrouded and highly loaded axial turbine configuration with 3-dimensionally shaped blades and cylindrical end walls has been tested in an axial turbine facility. Both, the blade design and the rim seal purge flow path are representative to modern high pressure gas turbines.
The unsteady flow field at the hub cavity exit region has been measured with the fast-response aerodynamic probe (FRAP) for two different rim seal purge flow rates. Furthermore, fast-response wall mounted pressure transducers have been installed inside the cavity. Unsteady full-annular CFD simulations have been employed in order to complement the experimental work.
The time-resolved pressure measurements inside the hub cavity reveal clear cavity modes which show a strong dependency on the injected amount of rim seal purge flow. The numerical predictions provide information on the origin of these modes and relate them to pronounced ingestion spots around the circumference.
The unsteady probe measurements at the rim seal interface show that the signature of the hub cavity induced modes migrates into the main annulus flow up to 30% blade span. Based on that, an aerodynamic loss mechanism has been found, showing that the benefit in loss reduction by decreasing the rim seal purge flow rate is weakened by the presence of turbine hub cavity modes.