In the present paper, the synergistic combination of inter-cooling with pulsed detonation combustion is analyzed concerning its contribution to NO x and CO 2 emissions. CO 2 is directly proportional to fuel burn and can, therefore, be reduced by improving specific fuel consumption and reducing engine weight and nacelle drag. A model predicting NO x generation per unit of fuel for pulsed detonation combustors, operating with jet-A fuel, is developed and integrated within Chalmers University’s gas turbine simulation tool GESTPAN. The model is constructed using CFD data obtained for different combustor inlet pressure, temperature and equivalence ratio levels. The NO x model supports the quantification of the trade-off between CO 2 and NO x emissions in a 2050 geared turbofan architecture incorporating intercooling and pulsed detonation combustion and operating at pressures and temperatures of interest in gas turbine technology for aero-engine civil applications.

Modern gas turbine endwall is operating in harsher conditions for the application of low NOx combustor. Non-axisymmetric endwall has been extensively studied for aerodynamic performance improvement, because endwall contouring can decrease the pressure gradient between the pressure side (PS) and the suction side (SS) in the blade passage. In addition to the influence of pressure gradient on aerodynamic losses, the vortical structures induced by pressure gradient are also the sources of high heat transfer regions in the passage. Consequently, thermal loads might be reduced by decreasing the pressure gradient thus weakening the strength of the secondary flows. In terms of engineering applications, distribution of thermal load is very important for the design of endwall cooling scheme, and it is necessary to take both aerodynamic and heat transfer performances into consideration for the endwall profile design. In this work, aero-thermal coupled design optimization of a turbine blade endwall was carried out. The endwall contour was obtained by multiplying heights of two curves in the streamwise and pitchwise directions. The streamwise curve was controlled by non-uniform B-spline (NUBS) and the pitchwise one was obtained by employing the sinusoidal function. The optimization method adopted in this research was the multi-objective genetic algorithm (MOGA) coupled with Kriging (KRG) model, which has been validated by benchmark functions. Numerical validation shows that static pressure coefficients on the blade surfaces and the Nusselt number ( Nu ) on the endwall agree well with the experimental results. The design variables were the endwall profile parameters, and the objective functions were maximizing total pressure recovery coefficient (ξ) at the blade outlet and minimizing the Nu on the endwall. Two optimal cases were selected from the Pareto front and analyzed in detail. It is indicated that the turbine blade aerodynamic performance can be improved while the heat transfer is restrained simultaneously. For the optimal Case I, mass flow-averaged ξ increases by 0.88%, and for Case II, area-averaged Nu reduces by about 7.78%.

As bypass-ratio in modern aero engines is continuously increasing over the last decades, the radial offset between low pressure compressor (LPC) and high pressure compressor (HPC), which needs to be overcome by the connecting s-shaped intermediate compressor duct (ICD), is getting higher. Due to performance and weight saving aspects the design of shorter and therefore more aggressive ducts has become an important research topic. In this paper an already aggressive design (with respect to current aero engines) of an ICD with integrated outlet guide vane (OGV) is used as a baseline for an aerodynamic optimization. The aim is to shorten the duct even further while maintaining it separation free. The optimization is broken down into two steps. In the first optimization-step the baseline design is shortened to a feasible extent while keeping weak aerodynamic restrictions. The resulting highly aggressive duct (intermediate design), which is shortened by 19 % in axial length with respect to the baseline, shows separation tendencies of low momentum fluid in the strut/hub region. For the second step, the length of the optimized duct design is frozen. By implementing new design features in the process of the optimizer, this optimization-step aims to eliminate separation and to reduce separation tendencies caused by the aggressive shortening. In particular, these features are: a nonaxisymmetric endwall contouring and parametrization of the strut and the OGV to allow for changes in lift and turning in both blade designs. By comparison of the three designs: Baseline, intermediate (separating flow) and final design, it can be shown, that it is possible to decrease length of the already aggressive baseline design even further, when adding a nonaxisymmetric endwall contouring and changes in blade shape of the strut and OGV. Flow separation can be eliminated while losses are kept low. With a more aggressive and therefore shorter duct the engine length and weight can be reduced. This in turn leads to lighter aircrafts, less fuel consumption and lower CO 2 and NO x emissions.

Ammonia is expected to be a hydrogen carrier that has potential as a carbon-free fuel. Ammonia is known as a nonignitable fuel, and it is not easy to hold ammonia flames under atmospheric conditions. A demonstration test with the aim of showing the potential of ammonia-fired power plants was conducted using a micro gas turbine. A 50-kW-class turbine system firing kerosene was selected as a base model. More than 40 kW of power generation was achieved by firing ammonia gas or a mixture of ammonia and methane by modifying the combustor, the fuel control device, and the gas turbine startup sequence. The prototype bifuel combustor is a swirl combustor employing a non-premixed flame and a decreased air flow rate near a gas fuel injector for flame holding. Ammonia combustion in the prototype bifuel combustor was enhanced by supplying hot combustion air and by modifying the air inlets. However, the exhaust gases from the ammonia flames had high NOx concentrations. NOx removal equipment using selective catalytic reduction can reduce NOx emission levels to below 10 ppm from more than 1000 ppm (converted value of NOx to 15% O 2 ) as already reported. However, downsizing of NOx removal equipment should be achieved for practical use. Therefore, a low NOx combustor was developed. As the first step of the development of the combustor, flame observation in the gas turbine combustor was tried. Although the observation area was limited, an inhomogeneous swirling orange flame of ammonia gas was observed. Then, a combustor test rig was prepared for a detailed observation of ammonia flame under various combustion conditions. The combustor test rig used a regenerative heat exchanger for heating combustion air, and it used an orifice for pressure drop instead of a turbine. Combustion air and cooling air were supplied from two air compressors. At startup of the combustor test rig, a spark plug was used to ignite non-premixed methane and air. After heating the regenerative heat exchanger, ammonia gas was supplied to the combustor instead of methane gas. The exhaust gases from the combustor were analyzed using FTIR (Fourier transform infrared spectroscopy) under various conditions, such as methane firing, methane–ammonia firing, and ammonia firing. Although there are several concepts for NOx reduction, a rich–lean combustion method was applied first for ammonia firing. The rich–lean combustor modified from the prototype bifuel combustor also could burn ammonia well in cases of both methane–ammonia firing and ammonia firing. The rich–lean combustor succeeded in reducing NOx emission from methane–ammonia combustion to half the value measured in the case of the prototype bifuel combustor.

Lean burn combustion is increasing its popularity in the aeronautical framework due to its potential in reducing drastically pollutant emissions (NO x and soot in particular). Its implementation, however, involves significant issues related to the increased amount of air dedicated to the combustion process, demanding the redesign of injection and cooling systems. A reduced coolant mass flow rate in conjunction with higher compressor discharge temperature negatively affect the cooling potential thus requiring the exploitation of efficient schemes such as effusion cooling. This work describes the experimental and numerical final validation of an aeronautical effusion-cooled lean-burn combustor. Full annular tests were carried out to measure temperature profiles and metal temperature distributions at different operating conditions of the ICAO cycle. Such an outcome was obtained also with an in-house developed CHT methodology (THERM3D). RANS simulations with the Flamelet Generated Manifold combustion model were performed to estimate aerothermal field and heat loads, while the coupling with a thermal conduction solver returns the most updated wall temperature. The heat sink within the perforation is treated with a 0D correlative model that calculates the heat pickup and the temperature rise of coolant. The results highlight an overall good capability of the proposed approach to estimate the metal temperature distribution at different operating conditions. It is also shown how more advanced scale-resolving simulations could significantly improve the prediction of turbulent mixing and heat loads.

Optimized operation of gas turbines is discussed for a fleet of eleven GE LM2500PE engines at a Statoil North Sea offshore field in Norway. Three engines are generator drivers and eight engines are compressor drivers. Several of the compressor drive engines are running at peak load (T5.4 control), hence production rate is limited by the available power from these engines. The majority of the engines discussed run continuously without redundancy, hence gas turbine uptime is critical for the field’s production and economy. The performance and operational experience with online water wash at high water-to-air ratio, as well as successful operation at longer maintenance intervals and higher average engine performance are described. This work is based on long-term operation with online washing, where operational data are collected and performance is analyzed over a 10-year period. Today, all engines are operated with 6-month intervals between maintenance stops, where offline crank wash is performed as well as other necessary maintenance and repairs. Online washing is performed daily between the maintenance stops at full load (i.e. normal operating load for the subject engine). To keep the engine as clean as possible and reduce degradation between maintenance stops, both an effective online water wash system and an effective air intake filter system are critical factors. The overall target is to maintain high engine performance, and extend the interval between maintenance stops through effective online washing. Water-to-air ratio is significantly increased compared to the OEM limit (OEM limit is 17 l/min which yields approx. 0.5% water-to-air ratio). Today the engines are operated at a water rate of 50 l/min (3 times the OEM limit) which yields a 1.4% water-to-air ratio. Such a high water-to-air ratio has been proven to be the key parameter for obtaining good online water wash effectiveness. Possible downsides of high water-to-air ratio have been thoroughly studied. The effect of optimized online water wash for the subject engines is longer intervals between maintenance stops, higher power availability, lower engine performance deterioration and reduced emissions (CO 2 and NO x ). The operating intervals are now extended to six months (4,000 hours), from initially two months (1,500 hours, early 1990s) followed by four months (3,000 hours, mid-2000s). Other installations operated as low as 750 hours between offline washes in the 1980s and 1990s. Of a total efficiency deterioration improvement of 6% over each 6-month operating period, the deterioration is reduced by an estimated 3% related to online water wash.

On two-shaft DLE (Dry Low Emissions) gas turbines compressor air is often bled off the combustor housing to control the combustor flame temperature for part load emissions management. The bleed air is typically directed into the exhaust ducting, bypassing the turbine, causing engine efficiency to decrease. A prototype bleed air recirculation system was fitted onto a compressor package in the field that redirects bleed flow from the engine exhaust to the air inlet to raise the air temperature entering the engine. This system is designed to reduce the amount of bleed air required at a given ambient operating condition, while not compromising emissions, to increase part load efficiency. The bleed air recirculation system was designed, built, installed, and tested on a Mars 16000S compressor package at a station in northern Pennsylvania. A control algorithm was developed and validated to control bleed air recirculation so that package durability was not impacted. Over an evaluation period of 12 months, no significant durability issues or degradation were noted. The on-site evaluation indicated that the engine’s heat rate is significantly improved by using bleed air recirculation. The heat rate was reduced by as much as 7% during testing at site with a strong dependency on the engine operating load and ambient temperature. As predicted the best case was demonstrated at 50% load and at colder ambient conditions. The NOx emissions were also improved with bleed air recirculation. In addition, using this system extended the range of low emissions operation to lower loads. No effects on compressor surge or sound emissions were detected.

This work is focused on investigations of Intercooled Recuperation configurations for aero engines. The investigated configurations were the Conventional Recuperation, where the recuperator was placed downstream the low-pressure turbine (LPT), the Alternative Recuperation, where the recuperator was mounted between the intermediate pressure turbine (IPT) and LPT, and the Staged Heat Recovery concept, in which two recuperators were mounted, one between IPT and LPT and the other downstream the LPT. These concepts were further assessed with the use of an additional combustor between IPT and LPT. All recuperator concepts that have been developed within the framework of various research projects by Aristotle University of Thessaloniki and MTU Aero engines AG were utilized, named as the NEWAC, CORN and STARTREC concepts. Additionally, a new recuperator design was introduced using a secondary fluid as heat transfer medium. The configurations thermodynamic assessment was focused on specific fuel consumption reduction, while the configurations effect on NOx emissions was also investigated. The results showed that the most promising recuperator concept, with respect to low TSFC values compared with a reference GTF engine of 2050 year technology level, is the recuperator with the secondary fluid heat transfer medium.

Micro gas turbines (MGT) used as combined heat and power (CHP) systems present interesting advantages over conventional CHP systems. For low power MGTs (1–3 kW el ), a single stage jet-stabilized combustor is well suited to reach low emissions despite high preheat temperatures. Although several jet stabilized combustors in this power range have already been presented in literature, the influence of single design parameters was not described yet. Therefore this paper presents a design study with different combustor geometries. Besides the number of nozzles, also the diameter of air and fuel nozzles as well as the length of the mixing zone is varied. Experiments are performed in an atmospheric combustor test rig with preheat temperatures of 730°C. Additionally RANS-CFD simulations for selected conditions are carried out. The results indicate an inverse correlation between the number of nozzles (diameter adjusted for constant inlet velocities) and the emissions, where an increase in the number of nozzles leads to lower emissions and a shorter reaction zone. Changing the air nozzle diameter and therefore the jet velocity creates a different trend. As the pressure loss decreases with smaller inlet velocities, the CO- and NO x -emission levels increase. An extension of the mixing section as well as of the diameter of the fuel nozzles have only minor effects for the examined geometries. The knowledge acquired in this work helps to evaluate the influence of design parameter changes on the behavior of small scale jet-stabilized combustors and allows an optimization of the combustor design regarding flame length and emissions.

This paper presents a numerical framework for characterizing fuel injection in modern combustors. The approach utilizes scaling analysis to describe the droplet evaporation in non-dimensional and fluid-independent terms. The results of the model are validated against published experimental data of isolated droplets evaporating at subcritical and near-critical conditions. The model is incorporated in a spray calculation framework and extended to the supercritical regime to assess the impact of different fluid-properties and evaporation models on temperature and fuel vapor distributions. The results suggest that in a non-convective environment the transient and quasi-steady evaporation rates vary exponentially with Lewis number. Furthermore, the results show fluid-independent behavior of the droplet evaporation, indicating that a single-component fluid can potentially be used as a modeling surrogate for jet fuel. The first-principles analysis demonstrates that classical evaporation models overestimate transient evaporation and underestimate quasi-steady evaporation, with discrepancies up to 70% at supercritical conditions. This is due to limitations in fuel-property description and the lack of non-isothermal droplet characterization at near-critical conditions. The temperature profiles are typically under-predicted and fuel vapor concentrations are over-predicted in standard spray calculations with subcritical evaporation models. As such the proposed framework breaks new ground in modeling of supercritical fuel injection. The improved quality in the predicted fuel concentration and temperature distribution can enable more accurate assessment of flame position, improving the estimation of combustion stability margins and NOx emissions. The model can be incorporated in commercial codes to guide the design of combustors operating at supercritical conditions.

Many sources of uncertainty exist when emissions are modelled for a gas turbine combustion system. They originate from uncertain inputs, boundary conditions, calibration, or lack of sufficient fidelity in the model. In this paper, a non-intrusive polynomial chaos expansion (NIPCE) method is coupled with a chemical reactor network (CRN) model using Python to rigorously quantify uncertainties of NO x emission in a premixed burner. The first objective of the uncertainty quantification (UQ) in this study is development of a global sensitivity analysis method based on NIPCE to capture aleatory uncertainty due to the variation of operating conditions and input parameters. The second objective is uncertainty analysis of Arrhenius parameters in the chemical kinetic mechanism to study the epistemic uncertainty in the modelling of NO x emission. A two-reactor CRN consisting of a perfectly stirred reactor (PSR) and a plug flow reactor (PFR) is constructed in this study using Cantera to model NO x for natural gas at the relevant operating conditions for a benchmark premixed burner. UQ is performed through the use of a number of packages in Python. The results of uncertainty and sensitivity analysis using NIPCE based on point collocation method (PCM) are then compared with the results of advanced Monte Carlo simulation (MCS). Surrogate models are also developed based on the NIPCE approach and compared with the forward model in Cantera to predict NO x emissions. The results show the capability of NIPCE approach for UQ using a limited number of evaluations to develop a UQ-enabled emission prediction tool for gas turbine combustion systems.

In the world economy more and more attention is paid to the environment protection. This has brought a requirement for reduction of harmful substances emission from the gas turbine engine combustors to the foreground scene. Several concepts of low-emission combustion of liquid fuels have been suggested to solve the problem of nitric oxide emission reduction. The authors consider combustion of lean homogenized (quick-mixed) fuel-air mixtures to be the most promising concept for a multi-mode combustion chamber. Based on the accumulated experience, the authors have formed some notion with respect to design peculiarities of low-emission combustors. Based on such general notions, an attempt has been made to create a model combustion chamber for decreasing harmful substances emission. A design for compact mixing modules has been worked out, as well as for a perforated flame tube. 3D computations have been carried out for the flow in the combustor compartment with 3 mini-modules, so to compare design and experimental data. In calculations the air entered the flame tube through a channel with a rectangular cross-section and, further, through swirlers of three burners (60% of air flow). Besides, the air came into the gap between the flame tube and casing through two side channels and, further, it got inside the flame tube through cooling system holes (40% of air flow). In parallel, tests have been carried out in similar combustor compartment, using standard fuels, measuring harmful substances emission at gas temperature (T 4 ) up to 1700 K. Data obtained testifies to essential reduction of nitric oxides in the experimental combustor being considered. Emission index NOx does not exceed value of 1 g/kg f in all the conditions investigated. Fuel efficiency is ≥ 99% for all the measurement regimes, except one, where it is 98%. Additionally, tests have been conducted, using bio fuel obtained from plant raw material. Research results have revealed problems of changeover to such type of fuel mixtures. Comparing test data with 3D simulation results, it can be noted that there, where computed value of the fuel combustion efficiency coincides with the measured one, NOx value also coincides. However, the emission index value is higher there, where the fuel combustion efficiency value obtained in computation is higher, i.e. where there are zones with higher temperature. The experimental results obtained have confirmed possibility of organizing low-emission combustion, as well as possibility of achieving the nitric oxide emission index level equal to 1 g/kg f at the combustor inlet temperature of 682K. It is evident that more detailed design study is required for transfer of the experimental technology to the working compartment of the combustion chamber. The achieved level of harmful substances emission, after improvement and implementation of technology, may allow meeting the strictest ICAO requirements and reducing the airport fees significantly.

Increasingly stringent regulations on emissions in the gas turbine industry require novel designs to minimize the environmental impact of oxides of nitrogen (NO x ). The development of advanced low-NO x technologies depends on accurate and reliable thermochemical mechanisms to achieve emissions targets. However, current combustion models have high levels of uncertainty in kinetic rates that, when propagated through calculations, yield significant variations in predictions. A recent study identified and optimized nine elementary reactions involved in CH formation to accurately capture its concentration and improve prompt-NO predictions. The current work quantifies the uncertainty on peak CH concentration and NO x emissions generated by these nine reaction rates only, when propagated through the San Diego mechanism. Various non-intrusive spectral methods are used to study atmospheric alkane-air flames. 1 st - and 2 nd -order total-order expansions and tensor-product expansions are compared against a reference Monte Carlo analysis to assess the ability of the different techniques to accurately quantify the effect of uncertainties on the quantities of interest. Sparse grids, subsets of the full tensor-product expansion, are shown to retain the advantages of tensor formulation compared to total-order expansions while requiring significantly fewer collocation points to develop a surrogate model. The high resolution per dimension can capture complex probability distributions witnessed in radical species concentrations. The uncertainty analysis of lean to rich flames demonstrated a high variability in NO x predictions reaching up to 400 % of nominal predictions. Wider concentration intervals were observed in rich conditions where prompt-NO x is the dominant contributor to emissions. The high variability and scale of uncertainty in NO x emissions originating from these nine elementary reactions demonstrate the need for future experiments and data assimilation to constrain current models to accurately capture CH for robust NO x emissions predictions.

One method frequently used to reduce NOx emissions is exhaust gas recirculation (EGR), where a portion of the exhaust gases, including NOx, is reintroduced into the combustion chamber. While a significant amount of research has been performed to understand the important fuel/NOx chemistry, more work is still necessary to improve the current understanding on this chemistry and to refine detailed kinetics models. To validate models beyond global kinetics data such as ignition delay time or flame speed, the formation of H 2 O was recorded using a laser absorption diagnostic during the oxidation of a mixture representing a simplistic natural gas (90% CH 4 /10% C 2 H 6 (mol.)). This mixture was studied at a fuel lean condition (equiv. ratio = 0.5) and at atmospheric pressure. Unlike in conventional fuel-air experiments, NO 2 was used as the oxidant to better elucidate the important, fundamental chemical kinetics by exaggerating the interaction between NOx and hydrocarbon-based species. Results showed a peculiar water formation profile, compared to a former study performed in similar conditions with O 2 as oxidant. In the presence of NO 2 , the formation of water occurs almost immediately before it reaches more or less rapidly (depending on the temperature) a plateau. Modern, detailed kinetics models predict the data with fair to good accuracy overall, while the GRI 3.0 mechanism is proven inadequate for reproducing CH 4 / C 2 H 6 and NO 2 interactions.

Liquefied petroleum gas (LPG) will be suitable for satisfying part of the growing global energy demands. The widespread utilization of LPG as a gas turbine fuel for power generation requires an advanced combustor that achieves dry low nitrogen oxides (NOx) combustion and flashback-resistant combustion. This paper describes the development of a “multi-cluster combustor” as an advanced dry low NOx and flashback-resistant combustion technology for dual gaseous fuels of natural gas and petroleum gas. The dual gaseous fuel capability will contribute to expanding fuel flexibility. The purpose of this paper is to evaluate the feasibility of the dual gaseous fueled combustion with the multi-cluster combustor with the same configuration. The combustor was tested in a single-can combustor test stand at medium pressure with both fuels. In the tests, natural gas consisted mainly of methane with a content of over 90 vol.%, and petroleum gas consisted almost entirely of propane. The test results showed that the combustor achieves dry low NOx combustion of both fuels within their stable ranges without flashback. This paper concluded from the test results that the multi-cluster combustor possesses the potential capability to achieve dry low NOx and flashback-resistant combustion of dual gaseous fuels of natural gas and petroleum gas. As the next step, further tests will be required with petroleum gas including butane and for high pressure conditions.

This work is motivated by an interest in understanding the fuel interchangeability of a fuel nozzle to operate under extreme lean operating conditions. A lean premixed, swirl-stabilized fuel nozzle designed with central pilot hub was used to test various fuel blends for combustion characteristics. Current gas turbine combustion technology primarily focuses on burning natural gas for industrial systems. However, interest in utilizing additional options due to environmental regulations as well as concerns about energy security have motivated interest in using fuel gases that have blends of Methane, Propane, H 2 , CO, CO 2 , and N 2 . For example, fuel blends of 35%/60% to 55%/35% of CH 4 /CO 2 are typically seen in Landfill gases. Syngas fuels are typically composed primarily of H 2 , CO, and N 2 . CH 4 /N 2 fuel blend mixtures can be derived from biomass gasification. Stringent emission requirements for gas turbines stipulate operating at extreme lean conditions, which can reduce NO x emissions. However, lean operating conditions pose the problem of potential blowout resulting in loss of performance and downtime. Therefore, it is important to understand the Lean Blowout (LBO) limits and involved mechanisms that lead to a blowout. While a significant amount of research has been performed to understand lean blowout limits and mechanisms for natural gas, research on LBO limits and mechanisms for fuel blends has only been concentrated on fuel blends of CH 4 and H 2 such as syngas. This paper studies the lean blowout limits with fuel blends CH 4 -C 3 H 8 , CH 4 -CO 2 , and CH 4 -N 2 and also their effect on the stability limits as the pilot fuel percentage was varied. Experimental results demonstrate that the addition of propane, nitrogen and carbon dioxide has minimal effect on the adiabatic flame temperature when the flame becomes unstable under lean operating conditions. On the other hand, the addition of diluent gas showed a potential blowout at higher adiabatic temperatures.

The Moderate or Intense Low-oxygen Dilution (MILD) combustion is characterized by low emission, stable combustion and low noise for various kinds of fuel. MILD combustion is a promising combustion technology for gas turbine. The model combustor composed of an optical quartz combustor liner, four jet nozzles and one pilot nozzle has been designed in this study. The four jet nozzles are equidistantly arranged in the combustor concentric circle and the high-speed jet flows from the nozzles will entrain amount of exhaust gas to make MILD combustion occur. The combustion characteristics of the model combustor under atmosphere pressure have been investigated through experiments and numerical simulations. The influence of equivalence ratio and jet velocity on flow pattern, combustion characteristics and exhaust emissions were investigated in detail, respectively. Laser Doppler velocity (LDV) was utilized to measure the speed of a series of points in the model combustor. The measurement results show that a central recirculation existed in the combustion chamber. As the jet velocity of the nozzles increases, the amount of entrained mass by the jet increases simultaneously, however, the central recirculation zone is similar in shape and size. The recirculation of the model combustor will remain self-similar when the jet velocity varies in the range. The calculation model and method were verified through comparing with experimentally LDV data and be used to optimize the model combustor. Planar laser-induced fluorescence of hydroxyl radical (OH-PLIF) approaches were adopted to investigate the flame structure, the reaction zone and the OH distribution. OH distribution of the paralleled and crossed sections in the model combustor were measured, the whole reaction zone have been analyzed. The results show that the OH distribution was uniform in whole combustor. The exhaust gas composition of the combustor was measured by the “TESTO 350” Exhaust Gas Analyzer. All measurements emission results were corrected to 15% O 2 in volume. Experimental results showed that NOx and CO emissions were less than 10 ppm@15%O 2 when the equivalence ratio ranges from 0.63 to 0.8.

A novel dry low-NOx gas turbine technology requires well balanced assessments since the early development phases. The weak knowledge of often conflicting aspects, such as operability and manufacturability, make any roadmap difficult to be drawn. The introduction of innovative manufacturing technologies such as the Direct Metal Laser Sintering (DMLS) process allows rapid manufacturing of components and test them in dedicated facilities to support real-time development of new products. The use of such a manufacturing process allows the adoption of designed experiments based development strategies, which are still uncommon at industrial level, due to the reduced time from drawings to test. The paper describes a reactive test campaign performed by BHGE in cooperation with University of Florence, aimed at the exploration of capabilities of different innovative burners in terms of pollutant emissions containment and blow-out margin. In particular, the test campaign has been conceived to provide a robust estimate of the effects of key geometrical parameters on principal burner performances. The flame stabilization mechanism of the investigated burners is based on the swirling flow generated by different setup of two internal channels: co-rotating and counter-rotating radial and axial swirlers. The effect of both the shape and the size of the internal air passages, as well as of the swirler characteristics, has been matter of investigation. Burners were tested in a single-cup test rig operated at moderate pressure conditions (up to 6bar), with two levels of preheated air temperature (300°C and 400°C). Each burner was equipped with two natural gas feeding lines representing the diffusion (pilot) and premixed (main) fuel supplies: both lines were regulated during tests to assess the effect of fuel split on emissions and to identify a stable low-NOx operating window, within which a lean blow-out test was performed. Dynamic pressure probes, were used to evaluate the onset of combustion instabilities. The burner development was supported by CFD investigations with the purpose to have a detailed understating of the flow-field and flame structure and to perform a preliminary screening to select the most promising solutions for the testing phase. The post process of the experimental results has allowed to correlate the main design parameters to burner performance variables discovering possible two-fold optimizations in terms of emissions and operability.

Modern gas turbines in power systems employ lean premixed combustion to lower flame temperature and thus achieve low NO x emissions. The fuel/air mixing process and its impacts on emissions are of paramount importance to combustor performance. In this study, the mixing process in a methane-fired model combustor was studied through an integrated experimental and numerical study. The experimental results show that at the dump location, the time-averaged fuel/air unmixedness is less than 10% over a wide range of testing conditions, demonstrating the good mixing performance of the specific premixer on the time-averaged level. A study of the effects of turbulent Schmidt number on the unmixedness prediction shows that for the complex flow field involved, it is challenging for Reynolds-Averaged Navier-Stokes (RANS) simulations with constant turbulent Schmidt number to accurately predict the mixing process throughout the combustor. Further analysis reveals that the production and scalar dissipation are the key physical processes controlling the fuel/air mixing. Finally, the NO x formation in this model combustor was analyzed and modelled through a flamelet-based approach, in which NO x formation is characterized through flame-front NO x and its post-flame formation rate obtained from one-dimensional laminar premixed flames. The effect of fuel/air unmixedness on NO x formation is accounted for through the presumed probability density functions (PDF) of mixture fraction. Results show that the measured NO x in the model combustor are bounded by the model predictions with the fuel/air unmixedness being 3% and 5% of the maximum unmixedness. In the context of RANS, the accuracy in NO x prediction depends on the unmixedness prediction which is sensitive to turbulent Schmidt number.

Micro gas turbine is one of the ideal prime movers for small-distributed energy systems. It can effectively reduce the emission of greenhouse gases and nitrogen oxides. Moreover, the use of micro gas turbines will contribute to burning fossil fuels in a much cleaner way. The staged combustion technology is the favorite way for low pollution combustion chamber such like. Therefore, the influence of the proportion of pilot fuel in the combustion chamber on pollutant emission deserves further study. The object of this research is the Double annular swirler (Das), which was applied to a 100 kW micro gas turbine combustion chamber. The combustion performance and emission characteristics under different Pilot Fuel Ratios (PFR) were obtained in prototype experimental system. Under the experimental conditions, Computational fluid dynamics (CFD) method was applied to research the reacting flow field and the formation of NO x in the combustion chamber and then analyze the influences of PFRs on combustion process. Experimental results show that the NO x emission of Das decreased at first and then increased with the augment of PFR. When PFR was near to 11%, the per unit NO x emission concentration reached its minimum. The numerical simulation agreed well with the experimental data. Further analysis of the simulation results indicate that there is a strong correlation between Φ local and NO x concentration. When it is lower than a certain value, the number of nitrogen oxides will be significantly reduced. The value has a lot to do with the inlet air temperature and the pressure of the combustion chamber under the design condition, and it needs to be confirmed by calculating the adiabatic temperature. Simultaneously, we also find that although the percentage of total air flowing into the combustor remains unchanged, the increase of PFR would reduce the airflow ratio in inner swirler. This implies that for some particular combustion chambers, special attention should be paid to the changes in air allocation caused by PFR.