133 research outputs found

    Optical Characterization of a Multipoint Lean Direct Injector for Gas Turbine Combustors: Velocity and Fuel Drop Size Measurements

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    Performance of a multipoint, lean direct injection (MP-LDI) strategy for low emission aero-propulsion systems has been tested in a Jet-A fueled, lean flame tube combustion rig. Operating conditions for the series of tests included inlet air temperatures between 672 and 828 K, pressures between 1034 and 1379 kPa and total equivalence ratios between 0.41 and 0.45, resulting in equilibrium flame temperatures approaching 1800 K. Ranges of operation were selected to represent the spectrum of subsonic and supersonic flight conditions projected for the next-generation of commercial aircraft. This document reports laser-based measurements of in situ fuel velocities and fuel drop sizes for the NASA 9-point LDI hardware arranged in a 3 3 square grid configuration. Data obtained represent a region of the flame tube combustor with optical access that extends 38.1-mm downstream of the fuel injection site. All data were obtained within reacting flows, without particle seeding. Two diagnostic methods were employed to evaluate the resulting flow path. Three-component velocity fields have been captured using phase Doppler interferometry (PDI), and two-component velocity distributions using planar particle image velocimetry (PIV). Data from these techniques have also offered insight into fuel drop size and distribution, fuel injector spray angle and pattern, turbulence intensity, degree of vaporization and extent of reaction. This research serves to characterize operation of the baseline NASA 9- point LDI strategy for potential use in future gas-turbine combustor applications. An additional motive is the compilation of a comprehensive database to facilitate understanding of combustor fuel injector aerodynamics and fuel vaporization processes, which in turn may be used to validate computational fluid dynamics codes, such as the National Combustor Code (NCC), among others

    Spray Flame and Exhaust Jet Characteristics of a Pressurized Swirl Combustor

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    This work describes an investigation of swirl-stabilized flames, created in a combustor featuring co-annular swirling airflows, under unenclosed, enclosed, and submerged conditions. A centrally-located fuel nozzle, which uses air-assist atomization, creates a methanol fuel spray. This approach provides great control over fuel spray properties in a compact geometry. Factors affecting the structure of the flames, including the effect of the central atomization air jet, are investigated using three-dimensional particle image velocimetry, direct imaging, and phase-Doppler particle analysis techniques. Exhaust jet temperatures are measured. The dynamic events affecting two-phase exhaust jets from the combustor under submerged conditions are examined using high-speed cinematography and sound spectrum analysis. It is found that the structures of the flames examined, which feature low overall equivalence ratios, are closely linked to the features of the air flowfield in the combustor. Swirl numbers of flows emerging from twisted-vane swirl assemblies are characterized. The structure of the flow is affected by the swirl configuration, but does not depend heavily on the Reynolds number. The central atomization air jet (with or without fuel) reshapes the recirculation region in the swirling flow and has a significant, controllable effect on the structure of the airflow and flame. The effect is the same for nonreacting and reacting flows. In one unique case, the central atomization air interacts with the swirling flow to create two recirculation regions and a lifted flame. The lifted flame is more compact than similar non-lifted flames. The twin-fluid atomization approach is shown to provide effective atomization over a wide range of operating conditions. The two-phase interaction of the exhaust jet is found to depend on the pressure drop over the exhaust nozzle. The dynamic behavior of the exhaust jet is buoyancy-driven at low pressure drops, and is affected by complex instability mechanisms at high pressure drops. Strouhal numbers of large-scale unstable events occurring in the two-phase flow are two orders of magnitude smaller than those associated with instabilities in single-phase flows. Evidence is presented, indicating that acoustic pressure waves in the exhaust jet may be involved in the generation of bubbles surrounding exhaust jets at high pressure drops

    The Role of CFD in Modern Jet Engine Combustor Design

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    Recent advances in the application of computational fluid dynamics (CFD) for turbulent combustion with the relevance for gas turbine jet engines are discussed. Large eddy simulation (LES) has emerged as a powerful approach to handle the highly turbulent, unsteady and thermochemically non-linear flows in the practical combustors, and it is a matter of time for the industry to replace the conventional Reynolds averaged Navier-Stokes (RANS) approach by LES as the main CFD tool for combustor research and development. Since combustion is a subgrid scale phenomenon in LES, appropriate modelling is required to describe the SGS combustion effects on the resolved scales. Among the various available models, the flamelet approach is seen to be a promising candidate for practical application because of its computational efficiency, robustness and accuracy. A revised flamelet formulation, FlaRe, is introduced to outline the general LES methodology for combustion modelling and then used for a range of test cases to demonstrate its capabilities for both laboratory burners and practical engine combustors. The LES results generally compare well with the experimental measurements showing that the important physical processes are captured in the simulations

    Application of Laser-based Diagnostics to a Prototype Gas Turbine Burner at Selected Pressures

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    The matured laser-diagnostic techniques of planar laser-induced fluorescence (PLIF) and particle image velocimetry (PIV) were applied to a prototype gas turbine burner operating on various fuels. The work was performed to provide verification of computational fluid dynamic (CFD) models of the combustion of atypical fuels in a gas turbine combustor. The burner was operated using methane and three synthesized fuels of interest- one with hydrogen as the principle component and two with a low heating value (15 MJ/m3). Experiments were performed at pressures from 1 to 9 bar, with the fuel/air mixture at both ambient (~ 300 K) and elevated temperature. The burner, which was supplied by Siemens Industrial Turbomachinery, is a down-scaled prototype of that used in the SGT-750 gas turbine. It is composed of three individual sectors that are arranged concentrically, a centermost pilot sector, intermediate sector and main sector. Each sector contributes a premixed fuel/air flow, while swirl elements in each sector promote flame stabilization and recirculation in the combustion region. There are dedicated fuel feeds allowing for localized setting of fuel/air mixture at each of the sectors. The central pilot sector of the burner was separable from the full burner assembly and was examined in detail. Information was generated regarding the use of syngas to fuel the burner. This information is intended to be used for the validation of CFD models of the experiments, including optimization of reduced chemical kinetic mechanisms for the specific fuels. Laminar flame speed measurements were made for several syngas fuel candidates from which the high-hydrogen syngas fuel was selected. Burner performance at the lean stability limit was examined using the fuels of interest. It was found that increasing the fuel/air ratio in the central pilot sector improved the lean limit onset of flame extinction up to the point that the central pilot extinguished. Optimization of the burner nitrogen oxides (NOx) emission by fuel partitioning among the three sectors was performed. The response in emission level with fuel/air ratio was not universal among the fuels tested. The largest portion of work in this thesis is the visualization of the burner combustion field by laser diagnostic methods. The flame shape was imaged by the PLIF of the OH radical distribution. PLIF imaging of the central pilot sector was recorded for atmospheric and elevated pressure for iterations of inlet air temperature, fuel type and equivalence ratio. When comparing the OH-LIF distribution for various fuels and pressures it was found that equivalence ratio had the greatest effect on the distribution of OH signal from the exit of the central pilot sector. Lean equivalence ratios showed a diffuse signal typical of the post combustion region. Near stoichiometric equivalence ratios yielded a distribution having a clearly defined inner edge indicating combustion occurring outside of the pilot sector. At rich equivalence ratios the OH signal was lifted away from the pilot burner exit. Comparison of OH-PLIF and chemiluminescence signal for methane combustion supported the characterization that the pilot sector efflux varied from post combustion to attached and then lifted flame in conjunction with the increase in equivalence ratio from lean to rich. OH-PLIF imaging was collected for staging of fuel to all three sectors of the burner at atmospheric pressure. The flow field in the combustion region produced by the full burner was visualized using PIV for each of fuels of interest, illustrating the recirculation zone. Finally the OH-LIF distribution was imaged for the combustion region of the entire burner at elevated pressure during operation at a single equivalence ratio with various dilutions of natural gas. There was little discernible change in flame shape as the pressure was changed from 3, 4.5 and 6 bar and energy content was changed from 30, 40 and 45 MJ/m^3 Wobbe index

    Laser diagnosis of gas turbine fuel sprays; scaling effects on NOx emissions and stability

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    This thesis first provided strategic recommendations for the research sponsor, Rolls- Royce plc (RR) and then applied optical diagnostics to measure aero gas turbine fuel spray properties in order to predict Oxides of Nitrogen (NOx) emissions and combustion instability. Analysis of the large civil aero engine sector suggested possible courses of action for RR to protect itself from short-term market volatilities and also prepare for three long term changes in strategic operating context: air traffic growth; tighter United Nations enforced aero engine combustion emissions legislation and entry of civil aviation into the European Union Emissions Trading Scheme. A collaborative game theoretic approach was explored during the pre-competitive, pre-technology, capability acquisition aero engine design phase on unproven future technologies to reduce R&D expenditures, development times and the costs of failure. Lean Prevapourised Premixed combustion demands excellent spray atomisation quality to sustain combustion efficiency, stability and to minimise pollutants. Post development of an improved procedure to calibrate laser signals, methodology to predict NOx and technique to optimise rig operating conditions that minimised fractional discrepancies in two-phase flow behaviour with corresponding engine conditions, this thesis applied quantitative Planar Laser Induced Fluorescence (PLIF) and Laser Sheet Dropsizing (LSD) to measure the fuel placement and dropsize distribution in the near nozzle regions of RR liquid-fuelled hybrid, airblast and pressure-swirl sprays. Measurements were made under non-combusting, low pressure conditions and results were processed to identify fuel injector designs that exhibited low emissions and high stability for the Affordable Near Term Low Emissions (ANTLE) and Instability Control of Low Emission Aero-Engine Combustors (ICLEAC) engine demonstrator programmes. Results also provided validation data and boundary conditions for spray computational codes. Research findings will improve RR core competencies in fuel injection research to accelerate the development and deployment of low emissions aero engine technology

    Investigation into Mixing and Combustion in an Optical, Lean, Premixed, Prevaporised Combustor.

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    Gaseous and particulate emissions from combustion devices are implicated in many atmospheric environmental pollution concerns. Lean, premixed, prevaporised combustion is widely regarded as the most practical technique for reducing these emissions from gas turbine combustors to levels which will not cause significant environmental impact. This technique has been proved to be capable of reducing emissions of oxides of nitrogen to ultra low levels. However, further understanding and development is necessary before LPP combustors can be reliably fitted in production gas turbines. Particular problems are flashback and autoignition in the premixer and achieving a stable, lean primary zone. This thesis details a comprehensive series of measurements made upon a realistic LPP gas turbine combustor. The measurements elucidate the important, fundamental, physical processes which govern the performance of LPP combustors whilst providing a challenging and complete data set for CFD model validation. These measurements include data on the premixer velocity field, the fuel droplet size and velocities distributions, the fuel concentration in the premixer and primary zone and the combustion temperature. This has been interpreted to provide useful information such as the location and rates of fuel-air mixing, the proportion of temporal to spatial fluctuation in fuel concentration, the premixer swirl number, the flame brush thickness and the effect on mixing and placement of fuel fraction boiling point. It has been found that for mixing multi-component fuels in a duct, the rate of mixing and physical placement will depend on the boiling fraction of the fuel. High boiling point fractions evaporate later, experience longer droplet trajectories and mix much slower when compared to lower boiling fractions.Ph

    Numerical and experimental investigation of a gasturbine model combustor with axial swirler

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    This thesis presents the simulations of and experiments with the so-called CECOST burner. The CECOST burner is a gas turbine model combustor equipped with an axial swirler and placed in a laboratory for the lean premixed turbulent combustion of a range of fuels at atmospheric pressure. The operation with natural gas, methane, hydrogen-enriched methane and syngas from black-liquor gasification was investigated experimentally and numerically. The CECOST burner is modular and adaptive, which enables the separate investigation of different parameters. A wide section of the CECOST swirler is manufactured from quartz to enable optical access. The flow field in the combustion chamber was measured by particle image velocimetry (PIV). High-speed imaging of the chemiluminescence signal in the bandwidth of hydroxyl radical relaxation was performed to document the intensity distribution for flames throughout the operating range of the CECOST burner. The lower (lean blow out) and upper (flashback) fuel-air equivalence ratio limits of stable operation were determined. The flashback of the flame upstream into the mixing section was captured by high-speed imaging. In the stable operating range, planar laser-induced fluorescence of hydroxyl radicals (OH-PLIF) was recorded to assess the turbulent flame structure.Reynolds-averaged Navier-Stokes (RANS) simulations were performed to calculate the steady isothermal flow in the CECOST burner for parameter studies. The measured flow data was used to validate the geometry and numerical discretisation of the computational model of the burner. Large eddy simulation (LES) was carried out to investigate the combusting flow with temporal resolution and high accuracy

    Design and Analysis of a Disk-oriented Engine Combustor

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    In a novel approach to gas-turbine power production, an engine was designed and analyzed to use both a single-stage centrifugal compressor and single-stage radial in- ow turbine configured back-to-back. This air path reduced the axial length of the engine up to 60%, providing additional modularity in a gas-turbine engine that could be used to improve mobility of ground-based power units or increase the survivability of aircraft through the use of distributive propulsion. This increased modularity was made possible by the use of a circumferential ow combustor that substantially decreased the axial length of the burner and negated the need to return compressor radial ow to the axial direction, as found in conventional combustion approaches. The Disk-Oriented Engine was designed to incorporate swirling inlet ow from a centrifugal compressor and exhaust directly into a radial in-ow turbine, while still maintaining the initial swirl pattern out of the compressor. The configuration of the combustion cavity was evaluated through computational fluid dynamics. An iterative design approach was used to achieve desired ow characteristics and combustion dynamics through geometry shaping and placement of air supply holes. The result of this design process was a computational combustor model that accepted swirling inlet ow, dispersed that air and fuel about a unique u-bend circumferential combustion cavity, and exhausted in the radial direction to feed a radial in-ow turbine. Sustained combustion was simulated at design conditions with a 3% total pressure loss in the combustor and a turbine inlet pattern factor of 0.24, indicating that such a design could operate as a gas-turbine engine, while reducing axial length up to 60% compared to traditional systems of similar size and performance. Computational results were compared to experimental tests on fuel-air swirl injectors, providing qualitative and quantitative insight into the stability of the flame anchoring system. From this design, a full-scale physical model of the Disk-Oriented Engine Combustor was designed and built for combustion analysis and characterization

    The experimental flowfield and thermal measurements in an experimental can-type gas turbine combustor

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    In this study, experimental data was collected in order to create a test case that can be used to validate computational fluid dynamics (CFD) simulations and the individual models used therein for gas turbine combustor applications. In many cases, the CFD results of gas turbine combustors do not correlate well with experimental results. For this reason, there is a requirement to test the simulation method used before CFD can successfully be used for combustor design. This test case encompasses all the features of a gas turbine combustor such as a swirler, primary, secondary and dilution holes as well as cooling rings. Experiments were performed on the same combustor geometry for both non-reacting and reacting flows. The non-reacting flow experiments consisted of stereoscopic particle image velocimetry (PIV) measurements performed at various planes in the three zones of the combustor. Data was collected on planes, both in line with the holes and in between the holes of each zone. For the reacting experiments, the temperatures on the outlet plane were measured using a thermocouple rake, thus a temperature contour plot on the outlet plane was produced. Further, the combustor can was modified with passive inserts, which were tested to determine their influence on the outlet temperature distribution during reacting runs. In this set-up, the outlet velocity profiles were also measured using a Pitot tube during both non-reacting and reacting flows. In addition to the outlet temperature distribution and velocity profiles, images of the flame patterns were captured, which showed the positions of flame tongues, fluctuating flames and steady flames. Carbon burn patterns on the walls of the combustor liner were also captured. From the data collected during the reacting runs, the pattern factor, profile factor, overall pressure loss and pressure loss factor were calculated. The non-reacting experiments performed using the PIV, produced three-dimensional velocity vector fields throughout the combustor. These experiments were performed at various flow rates, which gave an indication of which features of the combustor flow were affected by the flow rate. When comparing the individual PIV images alongside one another, the temporal nature of the combustor flow was also evident. The reacting experiments revealed a hot region of exhaust gas around the outer edge of the exhaust while there was a cooler region in the centre of the outlet flow. The PIV flowfield results revealed the reason for then hot outer ring-like region was due to the path the hot gasses would take. The hot combustor gas from the primary zone diverges outwards in the secondary zone then is further forced to the outside by the dilution recirculation zone. The hot flow then leaves the combustor along the wall while the cooler air from the jets leaves the combustor in the centre. The experiments performed produced a large variety of data that can be used to validate a number of aspects of combustor simulation using CFD. The non-reacting experimental data can be used to validate the turbulence models used and to evaluate how well the flow features were modelled or captured during the non-reacting stage of the combustor simulation process. The typical flow features such as jet penetration depths and the position and size of the recirculation regions are provided for effective comparison. The thermal results presented on the outlet plane of the combustor can be used for comparison with CFD results once combustion is modelled. CopyrightDissertation (MEng)--University of Pretoria, 2010.Mechanical and Aeronautical Engineeringunrestricte
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