Dissertations / Theses on the topic 'Gas Turbine Engine Combustion'

As a result of growing concerns about the carbon emissions associated with the combustion of conventional hydrocarbon fuels, hydrogen is gaining more attention as a clean alternative. The combustion of hydrogen in air produces no carbon emissions. However, hydrogen-air combustion does have the potential to produce oxides of nitrogen (NOx), which are harmful pollutants. The production of NOx can be significantly curbed using lean premixed combustion, wherein hydrogen and air are mixed at an equivalence ratio (the ratio of stoichiometric to actual air in the combustion process) significantly less than 1.0 prior to combustion. Hydrogen is a good candidate for use in lean premixed systems due to its very wide flammability range. The potential for the stable combustion of hydrogen at a wide range of equivalence ratios makes it particularly well-suited to application in gas turbines, where the equivalence ratio is likely to vary significantly over the operating range of the machine.

The strong lean combustion stability of hydrogen-air flames is due primarily to high reaction rates and the associated high turbulent burning velocities. While this is advantageous at low equivalence ratios, it presents a significant danger of flashbackâ the upstream propagation of the flame into the premixing deviceâ at higher equivalence ratios. An investigation has been conducted into the operation of a specific hydrogen-air premixer design in a gas turbine engine. Laboratory tests were first conducted to determine the upper stability limits of a single premixer. Tests were then carried out in which eighteen premixers and a custom-fabricated combustor liner were installed in a modified Pratt and Whitney Canada PT6A-20 turboprop engine. The tests examined the premixer and engine operability as a result of the modifications. A computer cycle analysis model was created to help analyze and predict the behavior of the modified engine and premixers. The model, which uses scaled component maps to predict off-design engine performance, was integral in the analysis of premixer flashback which limited the operation of the modified engine.
Master of Science

Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2003.
Includes bibliographical references (p. 71-72).
As part of the MIT micro-gas turbine engine project, the development of a hydrocarbon-fueled catalytic micro-combustion system is presented. A conventionally-machined catalytic flow reactor was built to simulate the micro-combustor and to better understand the catalytic combustion at micro-scale. In the conventionally-machined catalytic flow reactor, catalytic propane/air combustion was achieved over platinum. A 3-D finite element heat transfer model was also developed to assess the heat transfer characteristics of the catalytic micro-combustor. It has been concluded that catalytic combustion in the micro-combustor is limited by diffusion of fuel into the catalyst surface. To address this issue, a catalytic structure with larger surface area was suggested and tested. It was shown that the larger surface area catalyst increased the chemical efficiency. Design guidelines for the next generation catalytic micro-combustor are presented as well.
by Jhongwoo Peck.
S.M.

Transpiration cooling as potential replacement of multi-hole effusion cooling for gas turbine engines combustion liner application is investigated by comparing their cooling film effectiveness based on the mass transfer analogy (CFEM). Pressure sensitive paint was used to measure CFEM over PM surfaces which was found to be on average 40% higher than multi-hole effusion cooling. High porosity PM with low resistance to flow movement were found to offer uneven distribution of exiting coolant, with large amounts leaving the trailing edge, leading to lopsided CFEM. Design of anisotropic PM based on PM properties (porosity, permeability, and inertia coefficient) were investigated using numerical models to obtain more uniform CFEM. Heat transfer analysis of different PM showed that anisotropic samples offered better thermal protection over isotropic PM for the same porosity. Comparison between cooling film effectiveness obtained from temperatures CFET against CFEM revealed large differences in the predicted protection. This is attributed to the assumptions made to apply CFEM, nonetheless, CFEM remains a good proxy to study and improve transpiration cooling. A method for creating a CAD model of designed PM is proposed based on critical characteristics of transpiration cooling for future use in 3D printing manufacturing.

Hydrogen-powered engines have been gaining increasing interest due to the global concerns of the effects of hydrocarbon combustion on climate change. Gas turbines are suitable for operation on hydrogen fuel. This thesis reports the results of investigations of the special requirements of the fuel controller for a hydrogen gas turbine. In this investigation, a digital fuel controller for a hydrogen-fueled modified Pratt and Whitney PT6A-20 turboprop engine was successfully designed and implemented. Included in the design are safety measures to protect the operating personnel and the engine. A redundant fuel control is part of the final design to provide a second method of managing the engine should there be a malfunction in any part of the primary controller.

Parallel to this study, an investigation of the existing hydrogen combustor design was performed to analyze the upper stability limits that were restricting the operability of the engine. The upstream propagation of the flame into the premixer, more commonly known as a flashback, routinely occurred at 150 shaft horsepower during engine testing. The procedures for protecting the engine from a flashback were automated within the fuel controller, significantly reducing the response time from the previous (manual) method. Additionally, protection measures were added to ensure the inter-turbine temperature of the engine did not exceed published limits. Automatic engine starting and shutdown procedures were also added to the control logic, minimizing the effort needed by the operator. The tested performance of the engine with each of the control functions demonstrated the capability of the controller.

Methods to generate an engine-specific fuel control map were also studied. The control map would not only takes into account the operability limits of the engine, but also the stability limits of the premixing devices. Such a map is integral in the complete design of the engine fuel controller.
Master of Science

This thesis describes an experimental program covering the development of certain simple combustion chamber modifications to alleviate smoke emissions from the Allison T56 turboprop engines operated by the Royal Australian Air Force. The work includes a literature survey, smoke emission tests on two variants of the T56 engine, flow visualisation studies of the combustion system in a water tunnel and combustion rig tests of a standard combustor and four possible modifications. The rig tests showed that reductions in smoke emissions of 80% were possible by simple modifications that reduced the primary zone equivalence ratio and improved mixing in that zone.

Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2018
Cataloged from PDF version of thesis.
Includes bibliographical references (pages 47-50).
Aviation NO[subscript x] emissions have an impact on air quality and climate change, where the latter is magnified due to the higher sensitivity of the upper troposphere and lower stratosphere. In the aviation industry, efforts to increase the efficiency of propulsion systems are giving rise to higher overall pressure ratios which results in higher NO[subscript x] emissions due to increased combustion temperatures. This thesis identifies that the trend towards smaller engine cores (gas generators) that are power dense and contribute little to the thrust output presents new opportunities for emissions control that were previously unthinkable when the core exhaust stream contributed significant thrust. This thesis proposes and assesses selective catalytic reduction (SCR), which is a post-combustion emissions control method used in ground-based sources such as power generation and heavy-duty diesel engines, for use in aero-gas turbines.
The SCR system increases aircraft weight and introduces a pressure drop in the core stream. The effects of these are evaluated using representative engine cycle models provided by a major aero-gas turbine manufacturer. This thesis finds that employing an ammonia-based SCR can achieve close to 95% reduction in NO[subscript x] emissions for ~0.4% increase in block fuel burn. The large size of the catalyst needs to be housed in the body of the aircraft and hence would be suitable for future designs where the engine core is also within the fuselage, such as would be possible with turbo-electric or hybrid-electric designs. The performance of the post-combustion emissions control is shown to improve for smaller core engines in new aircraft in the NASA N+3 time-line (2030-2035), suggesting the potential to further decrease the cost of the ~95% NO[subscript x] reduction to below ~0.4% fuel burn.
Using a global chemistry and transport model (GEOS-Chem) this thesis estimates that using ultra-low sulfur (<15 ppm fuel sulfur content) in tandem with post-combustion emissions control results in a ~92% reduction in annual average population exposure to PM₂.₅ and a ~95% reduction in population exposure to ozone. This averts approximately 93% of the air pollution impact of aviation.
by Prakash Prashanth.
S.M.
S.M. Massachusetts Institute of Technology, Department of Aeronautics and Astronautics

The current demands for higher performance in gas turbine engines can be reached by raising combustion temperatures to increase thermal efficiency. Hot combustion temperatures create a harsh environment which leads to the consideration of the durability of the combustor and turbine sections. Improvements in durability can be achieved through understanding the interactions between the combustor and turbine. The flow field at a combustor exit shows non-uniformities in pressure, temperature, and velocity in the pitch and radial directions. This inlet profile to the turbine can have a considerable effect on the development of the secondary flows through the vane passage. This thesis presents a computational study of the flow field generated in a non-reacting gas turbine combustor and how that flow field convects through the downstream stator vane. Specifically, the effect that the combustor flow field had on the secondary flow pattern in the turbine was studied. Data from a modern gas turbine engine manufacturer was used to design a realistic, low speed, large scale combustor test section. This thesis presents the results of computational simulations done in parallel with experimental simulations of the combustor flow field. In comparisons of computational predictions with experimental data, reasonable agreement of the mean flow and general trends were found for the case without dilution jets. The computational predictions of the combustor flow with dilution jets indicated that the turbulence models under-predicted jet mixing. The combustor exit profiles showed non-uniformities both radially and circumferentially, which were strongly dependent on dilution and cooling slot injection. The development of the secondary flow field in the turbine was highly dependent on the incoming total pressure profile. For a case with a uniform inlet pressure in the near-wall region no leading edge vortex was formed. The endwall heat transfer was found to also depend strongly on the secondary flow field, and therefore on the incoming pressure profile from the combustor.
Master of Science

Conventional gas turbine combustors operate with a designed drop in pressure over the length of the device. This is desired in order to encourage mixing within the combustor. Compared to this, pulse pressure gain combustors are an alternative to the conventional combustor that produces an increase in static pressure between the inlet and exhaust of the device. The removal of the combustor pressure loss increases the efficiency of the combustion process by increasing the amount of work produced. Many types of pulsed pressure gain combustors exist. Of these, the valveless pulse combustor is the simplest featuring no moving parts. Whilst some research has been conducted into investigating the performance and workings of a pulse combustor, little has been conducted with the view of cooling the combustor. This has been the focus for the research contained herein. The research has focussed on establishing an understanding of the heat transfer characteristics within a pulse combustor tailpipe. This has involved experimental, analytical and computational research on a pulse combustor as well as on a cold-flow model of a pulse combustor tailpipe. This has enabled a study into the feasibility of cooling a pulse combustor to be conducted. The research has found that for conditions where the unsteady velocity amplitude within the cold-flow model of the pulse combustor tailpipe exceeds the mean velocity, an enhancement to the heat transfer coefficient is measured compared to the value expected in a similar non-oscillating flow. When there is no enhancement to the heat transfer coefficient, the cyclic variation of the unsteady heat flux follows the variation of the unsteady pressure within the device. However, at times of enhancement, the instantaneous heat flux structure shows a large deviation from the structure of the pressure field driving the oscillations. This change is shown to be caused by the reversal in the near-wall velocity and may indicate a mechanism for the enhancement in the mean heat flux. The cooling feasibility study showed that with further investigation, it may be possible to cool a pulse combustor within a gas turbine engine.

Thesis (Ph.D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2000.
"February 2000."
Includes bibliographical references (p. 203-211).
As part of an effort to develop a microfabricated gas turbine engine capable of providing 10-50 Watts of electrical power in a package less than one cubic centimeter in volume, this thesis presents the design, fabrication, packaging and testing of the first combustion system for a silicon micro heat engine. The design and operation of a microcombustor is fundamentally limited by the chemical reaction times of the fuel, by silicon material and fabrication constraints, and by the inherently non-adiabatic nature of the operating space. This differs from the design of a modern macro combustion system that is typically driven by emissions, stability, durability and pattern factor requirements. The combustor developed herein is shown to operate at a power density level that is at least an order of magnitude higher than that of any other power-MEMS device (2000 MW/m 3), and establishes the viability of using high power density, silicon-based combustion systems for heat engine applications at the micro-scale. This thesis presents the development of two specific devices - the first device is a 3-wafer level microcombustor that established the viability of non-premixed hydrogen-air combustion in a volume as small as 0.066 cm 3, and within the structural constraints of silicon; the second device is known as the engine "static-structure", and integrated the 3-stack microcombustor with the other non-rotating components of the engine. Fabricated by aligned fusion bonding of 6 silicon wafers, the static structure measures 2.1 cm x 2.1 cm x 0.38 cm, and was largely fabricated by deep reactive ion etching through a total thickness of 3,800 pm. Packaged with a set of fuel plenums, pressure ports, fuel injectors, igniters, fluidic interconnects, and compressor and turbine static airfoils, this structure is the first demonstration of the complete hot flow path of a multi-level microengine. The 0.195 cm 3 combustion chamber has been tested for several tens of hours and is shown to sustain stable hydrogen combustion with exit gas temperatures above 1600K and combustor efficiencies as high as 95%. The structure also serves as the first experimental demonstration of hydrocarbon microcombustion with power density levels of 500 MW/m 3 and 140 MW/m 3 for ethylene-air and propane-air combustion, respectively. In addition to the development of the two combustion devices, this thesis also presents simple analytical models to identify and explain the primary drivers of combustion phenomena at the micro-scale. The chemical efficiency of the combustor is shown to have a strong correlation with the Damkohler number in the chamber, and asymptotes to unity for sufficiently large values of Da. The maximum power density of the combustor is also shown to be primarily limited by the structural and fabrication constraints of the material. Overall, this thesis synthesizes experimental and computational results to propose a simple design methodology for microcombustion devices, and to present design recommendations for future microcombustor development. Combined with parallel efforts to develop thin-film igniters and temperature sensors for the engine, it serves as the first experimental demonstration of the design, fabrication, packaging and operation of a silicon-based combustion system for power generation applications at the micro-scale.
by Amitav Mehra.
Ph.D.

Current thinking on the best propulsion system for a next-generation supersonic cruising (Mach 2 to Mach 4) aircraft is a mixed-flow turbofan engine with afterburner. This study investigates the performance increase of a turbofan engine through the use of isothermal combustion inside the high-pressure turbine (High-Pressure Turburner, HPTB) as an alternative form of thrust augmentation. A cycle analysis computer program is developed for accurate prediction of the engine performance and a supersonic transport cruising at Mach 2 at 60,000 ft is used to demonstrate the merit of using a turburner. When assuming no increase in turbine cooling flow is needed, the engine with HPTB could provide either 7.7% increase in cruise range or a 41% reduction in engine mass flow when compared to a traditional turbofan engine providing the sane thrust. If the required cooling flow in the turbine is almost doubled, the new engine with HPTB could still provide a 4.6% increase in range or 33% reduction in engine mass flow. In fact, the results also show that the degradation of engine performance because of increased cooling flow in a turburner is less than half of the degradation of engine performance because of increased cooling flow in a regular turbine. Therefore, a turbofan engine with HPTB will still easily out-perform a traditional turbofan when even more cooling than currently assumed is introduced. Closer examination of the simulation results in off-design regimes also shows that the new engine not only satisfies the thrust and efficiency requirement at the design cruise point, but also provides enough thrust and comparable or better efficiency in all other flight regimes such as transonic acceleration and take-off. Another finding is that the off-design bypass ratio of the new engine increases slower than a regular turbofan as the aircraft flies higher and faster. This behavior enables the new engine to maintain higher thrust over a larger flight envelope, crucial in developing faster air-breathing aircraft for the future. As a result, an engine with HPTB provides significant benefit both at the design point and in the off-design regimes, allowing smaller and more efficient engines for supersonic aircraft to be realized.
Ph. D.

Current thinking on the best propulsion system for a next-generation supersonic cruising (Mach 2 to Mach 4) aircraft is a mixed-flow turbofan engine with afterburner. This study investigates the performance increase of a turbofan engine through the use of isothermal combustion inside the high-pressure turbine (High-Pressure Turburner, HPTB) as an alternative form of thrust augmentation. A cycle analysis computer program is developed for accurate prediction of the engine performance and a supersonic transport cruising at Mach 2 at 60,000 ft is used to demonstrate the merit of using a turburner. When assuming no increase in turbine cooling flow is needed, the engine with HPTB could provide either 7.7% increase in cruise range or a 41% reduction in engine mass flow when compared to a traditional turbofan engine providing the sane thrust. If the required cooling flow in the turbine is almost doubled, the new engine with HPTB could still provide a 4.6% increase in range or 33% reduction in engine mass flow. In fact, the results also show that the degradation of engine performance because of increased cooling flow in a turburner is less than half of the degradation of engine performance because of increased cooling flow in a regular turbine. Therefore, a turbofan engine with HPTB will still easily out-perform a traditional turbofan when even more cooling than currently assumed is introduced. Closer examination of the simulation results in off-design regimes also shows that the new engine not only satisfies the thrust and efficiency requirement at the design cruise point, but also provides enough thrust and comparable or better efficiency in all other flight regimes such as transonic acceleration and take-off. Another finding is that the off-design bypass ratio of the new engine increases slower than a regular turbofan as the aircraft flies higher and faster. This behavior enables the new engine to maintain higher thrust over a larger flight envelope, crucial in developing faster air-breathing aircraft for the future. As a result, an engine with HPTB provides significant benefit both at the design point and in the off-design regimes, allowing smaller and more efficient engines for supersonic aircraft to be realized.
Ph. D.

An experimental facility was designed and benchmarked that could simulate the non-uniformities in the flow and thermal fields exiting real gas-turbine combustors. The design of the combustor simulator required analyses of the flow paths within a real combustor in a gas turbine engine. Modifications were made to an existing wind tunnel facility to allow for the installation of the combustor simulator. The overall performance of the simulator was then benchmarked through measurements of velocity, pressure, temperature, and turbulence using a straight exit test section to provide a baseline set of data. Comparisons of the measured quantities were made between two test cases that included a flow field with and without dilution flow.One of the major findings from this study was that the total pressure profiles exiting the combustor simulator in the near-wall region were different from a turbulent boundary layer. This is significant since many studies consider a turbulent boundary layer as the inlet condition to the turbine. Turbulent integral length scales were found to scale well with the dilution hole diameters and no dominant frequencies were observed in the streamwise velocity energy spectra. Dilution flow resulted in an increase in turbulence levels and mixing causing a reduction in the variation of total pressure and velocity. Adiabatic effectiveness levels were significantly reduced for the case with dilution flow in both the near combustor exit region and along the axial length of the straight exit test section.
Master of Science

This work assesses the feasibility of designing and implementing the wave rotor (WR), the pulse detonation engine (PDE) and the internal combustion wave rotor (ICWR) as part of novel Brayton cycles able to reduce emissions of future aircraft. The design and evaluation processes are performed using the simplified analytical solution of the devices as well as 1D-CFD models. A code based on the finite volume method is built to predict the position and dimensions of the slots for the WR and ICWR. The mass and momentum equations are coupled through a modified SIMPLE algorithm to model compressible flow. The code includes a novel tracking technique to ensure the global mass balance. A code based on the method of characteristics is built to predict the profiles of temperature, pressure and velocity at the discharge of the PDE and the effect of the PDEs array when it operates as combustion chamber of gas turbines. The detonation is modelled by using the NASA-CEA code as a subroutine whilst the method of characteristics incorporates a model to capture the throttling and non-throttling conditions obtained at the PDE's open end during the transient process. A medium-sized engine for business jets is selected to perform the evaluation that includes parameters such as specific thrust, specific fuel consumption and efficiency of energy conversion. The ICWR offers the best performance followed by the PDE; both options operate with a low specific fuel consumption and higher specific thrust. The detonation in an ICWR does not require an external source of energy, but the PDE array designed is simple. The WR produced an increase in the turbine performance, but not as high as the other two devices. These results enable the statement that a pressure rise combustion process behaves better than pressure exchangers for this size of gas turbine. Further attention must be given to the NOx emission, since the detonation process is able to cause temperatures above 2000 K while dilution air could be an important source of oxygen.

This body of work investigated the influence of alternative jet fuel properties on aviation gas turbine performance at threshold combustor operating conditions. It focused on altitude blowout performance and was in part motivated by results that were encountered during an aviation industry evaluation of synthetic kerosene that complied with the Jet A-1 specification, but differed from the fuel that was used as a reference in terms of some significant properties. As a consequence the relative impact of physical properties and reaction chemistry properties were of primary interest in this study. The thesis considered the potential to blend a range of different alternative jet fuel formulations which exhibited independent variations in properties relating to evaporation and reaction behaviour whilst still conforming to legislated physical fuel specifications. It further explored the potential for said variations having a detectable and significant influence on the simulated high altitude extinction behaviour in a representative aviation gas turbine combustor. Based on the findings, appropriate metrics were suggested for scientifically quantifying the appropriate properties and conclusions were drawn about the potential impact of alternative jet fuel properties on blowout performance. These subjects were addressed primarily through the theoretical analyses of targeted experimental programmes. The experimental design adopted a novel approach of formulating eight test fuels to reflect real-world alternative fuel compositions while still enabling a targeted evaluation of the influences of both physical and chemical reaction properties. A detailed characterisation was performed of the test fuels' physical and reaction properties. The extinction and spray behaviours of the fuels were then evaluated in a laboratory scale combustor featuring dual-swirl geometry and a single prefilming airblast atomiser. The various experimental data sets were interpreted within the context of a theoretical model analysis. In doing so the relative performance of alternative jet fuel formulations under laboratory burner conditions were translated to predict relative real world altitude performance. This approach was validated against aforementioned industry evaluation results and demonstrated to be consistent. A technically defensible explanation was provided for the previously unexplored anomalous altitude extinction results that were observed during the industry evaluation of synthetic jet fuel. A conclusive case was made for the extinction limit differences having been caused by the relative differences in chemical ignition delays of the fuels. The probability of volatility (distillation profile) and fuel physical properties playing a significant role in the impaired altitude performance was discredited. Evaporation-controlled combustion efficiency was, however, shown to become a significant factor at low air mass flow rates or when the fuel evaporation is compromised. The influence of flame speed and chemical ignition delays were investigated. Laminar flame speed was shown not to correlate with LBO, discrediting its use as a proxy for reaction rate. The study showed a correlation between the lean blowout behaviour of jet fuels and the ignition delays associated with their derived cetane numbers. Additionally, there was substantive support indicating that an even stronger correlation could be obtained by operating the IQT™ device that is used to measure these delays at an elevated temperature. The thesis makes a contribution towards the development of both technical understanding and practical tools for evaluating the potential operating limits of alternative jet fuel formulations.

Current and future legislation demands ever decreasing levels of pollution from gas turbine engines, and with combustor performance playing a critical role in resultant emissions, a need exists to develop a greater appreciation of the fundamental causes of unsteadiness. Particle Image Velocimetry (PIV) provides a platform to enable such investigations. This thesis presents the development of PIV measurement methodologies for highly turbulent flows. An appraisal of these techniques applied to gas turbine combustors is then given, finally allowing a description of the increased understanding of the underlying fluid dynamic processes within combustors to be provided. Through the development of best practice optimisation procedures and correction techniques for the effects of sub-grid filtering, high quality PN data has been obtained. Time average statistical data at high spatial resolution has been collected and presented for generic and actual combustor geometry providing detailed validation of the turbulence correction methods developed, validation data for computational studies, and increased understanding of flow mechanisms. These data include information not previously available such as turbulent length scales. Methodologies developed for the analysis of instantaneous PIV data have also allowed the identification of transient flow structures not seen previously because they are invisible in the time average. Application of a new `PDF conditioning' technique has aided the explanation of calculated correlation functions: for example, bimodal primary zone recirculation behaviour and jet misalignments were explained using these techniques. Decomposition of the velocity fields has also identified structures present such as jet shear layer vortices, and through-port swirling motion. All of these phenomena are potentially degrading to combustor performance and may result in flame instability, incomplete combustion, increased noise and increased emissions.

Combustion instability represents a significant problem in the application of low emission lean premixed combustion for gas turbines and has become one of the primary concerns in modern gas turbine industry. Effusion cooling has become common practice in gas turbine combustors and when calibrated, perforated combustor liners are able to attenuate combustion instability within a wide frequency range. However, the acoustic attenuation effect of perforated liner absorbers varies with a considerable number of flow and geometry influencing factors. The traditional approach of designing perforated combustor liners relies heavily on expensive and lengthy trial-and-error experimental practice. Computational fluid dynamics (CFD), especially large eddy simulation (LES) method has gained recognition as a viable tool for the simulation of unsteady flows and the phenomenon of combustion instability in gas turbine combustors. However, detailed resolution of the many small scale features, such as effusion cooling holes, is computationally very expensive and restricts the routine simulation of detailed engine geometries. In this thesis, a novel homogenous porous media model is proposed for the simulation of acoustic attenuation effect of gas turbine perforated liners. The model is validated against a number of well-acknowledged experiments and is shown to be able to predict acoustic attenuation properties of gas turbine liners both in the linear and non-linear absorption regimes and also the effect of bias flow, grazing flow and the temperature of flow on the acoustic properties of the liners. The model is applied to a large eddy simulation of a lab-scale premixed combustor "PRECCINSTA" and is demonstrated to successfully represent noise attenuation effects of perforated liner absorbers in both cold flow and reacting flow conditions. This model is able to provide a significant reduction in the overall computational time in comparison to directly resolved geometries, and can be applied as such a viable option for routine engineering simulation of perforated combustor liners.

Biomass is one of the main natural resources in Sweden.Increased utilisation of biomass for energy purposes incombined heat and power (CHP) plants can help the country meetits nuclear phase-out commitment. The present low-CO2 emissioncharacteristics of the Swedish electricity production system(governed by hydropower and nuclear power) can be retained onlyby expansion of biofuels in the CHP sector. Domestic Swedishbiomass resources are vast and renewable, but not infinite.They should be utilised as efficiently as possible in order tomeet the conditions for sustainability in the future.Application of efficient power generation cycles at low cost isessential for meeting this challenge. This applies also tomunicipal solid waste (MSW) incineration with energyextraction, which is to be preferred to landfilling.

Modern gas turbines and internal combustion engines firedwith natural gas have comparatively low installation costs,good efficiency characteristics and show reliable performancein power applications. Environmental and source-of-supplyfactors place natural gas at a disadvantage as compared tobiofuels. However, from a rational perspective, the use ofnatural gas (being the least polluting fossil fuel) togetherwith biofuels contributes to a diverse and more secure resourcemix. The question then arises if both these fuels can beutilised more efficiently if they are employed at the samelocation, in one combined cycle unit.

The work presented herein concentrates on the hybriddual-fuel combined cycle concept in cold-condensing and CHPmode, with a biofuel-fired bottoming steam cycle and naturalgas fired topping gas turbine or engine. Higher electricalefficiency attributable to both fuels is sought, while keepingthe impact on environment at a low level and incorporating onlyproven technology with standard components. The study attemptsto perform a generalized and systematic evaluation of thethermodynamic advantages of various hybrid configurations withthe help of computer simulations, comparing the efficiencyresults to clearly defined reference values.

Results show that the electrical efficiency of hybridconfigurations rises with up to 3-5 %-points in cold-condensingmode (up to 3 %-points in CHP mode), compared to the sum of twosingle-fuel reference units at the relevant scales, dependingon type of arrangement and type of bottoming fuel. Electricalefficiency of utilisation of the bottoming fuel (biomass orMSW) within the overall hybrid configuration can increase withup to 8-10 %-points, if all benefits from the thermalintegration are assigned to the bottoming cycle and effects ofscale on the reference electrical efficiency are accounted for.All fully-fired (windbox) configurations show advantages of upto 4 %-points in total efficiency in CHP mode with districtheating output, when flue gas condensation is applied. Theadvantages of parallel-powered configurations in terms of totalefficiency in CHP mode are only marginal. Emissions offossil-based CO2 can be reduced with 20 to 40 kg CO2/MWhel incold-condensing mode and with 5-8 kg CO2 per MWh total outputin CHP mode at the optimum performance points.

Keywords: Biomass, Municipal Solid Waste (MSW), Natural Gas,Simulation, Hybrid, Combined Cycle, Gas Turbine, InternalCombustion Engine, Utilization, Electrical Efficiency, TotalEfficiency, CHP.

Power generation from biomass based renewable energy technologies is a promising option in retrofitting our dependence in conventional power generation processes. The development of any society is not possible without sustainable energy and access to energy creates that environment that allows the world to thrive. Electricity access especially in developing regions of the world is of particular interest. This work provides results on electricity efficiency, the economic feasibility and environmental impact of biomass based power technologies in small scale setting using Aspen Plus software. The power generation processes analysed on standalone basis include - micro gas turbine, gas turbine, steam turbine, Stirling engine and internal combustion engine. Some of the processes are optimized in the design to suit the specific climate and available wood waste stream in Nigeria is considered in this work. Simulation results indicate that gas engines power technologies gave a better electric performance of more than 30% with its integration with biomass gasification technology in production of fuel gas. The stirling engine power technology shows a good prospect despite its yet to be commercial status. The modification of the engine (removal regenerator) gives a better electric efficiency. Also result shows that internal combustion engine process emits more of nitric oxides compared to other technologies which create doubts over its environmental compatibility. Economic studies show that for small scale power generation, internal combustion engines and stirling engines are economic feasible. Also, steam turbine and gas turbine illustrate why they are mostly applied in medium/large scale biomass power generation specially recommended to regions where more biomass resource are produced. The micro gas turbine power technology can also be applied in small scale despite its high total investment capital. Furthermore, the study shows that about from 1.8 million tonnes per year of saw dust (wood waste) produced from lumber industries in Nigeria, about 1.3 TWh of electricity can be generated from 1000 MW power plant. Power generation via the utilization of biomass prove to be a possible path to Nigeria’s economic, social and environmental sustainability but the extent to which this can achieved is strongly dependent institutional framework, investment, incentives and information policies.
Program: Masterutbildning i energi- och materialåtervinning

A comparison study between axial and radial swirler performance in a gas turbine can combustor was conducted by investigating the correlation between combustor flow field geometry and convective heat transfer at cold flow conditions for Reynolds numbers of 50,000 and 80,000. Flow velocities were measured using Particle Image Velocimetry (PIV) along the center axial plane and radial cross sections of the flow. It was observed that both swirlers produced a strong rotating flow with a reverse flow core. The axial swirler induced larger recirculation zones at both the backside wall and the central area as the flow exits the swirler, and created a much more uniform rotational velocity distribution. The radial swirler however, produced greater rotational velocity as well as a thicker and higher velocity reverse flow core. Wall heat transfer and temperature measurements were also taken. Peak heat transfer regions directly correspond to the location of the flow as it exits each swirler and impinges on the combustor liner wall. Convective heat transfer was also measured along the liner wall of a gas turbine annular combustor fitted with radial swirlers for Reynolds numbers 210000, 420000, and 840000. The impingement location of the flow exiting from the radial swirler resulted in peak heat transfer regions along the concave wall of the annular combustor. The convex side showed peak heat transfer regions above and below the impingement area. This behavior is due to the recirculation zones caused by the interaction between the swirlers inside the annulus.
Master of Science

The highly competitive gas turbine industry has been motivated by consumer demands for higher power-to-weight ratios, increased thermal efficiencies, and reliability while maintaining affordability. In its continual quest, the industry must continually try to raise the turbine inlet temperature, which according to the well-known Brayton cycle is key to higher engine efficiencies. The desire for increased turbine inlet temperatures creates an extremely harsh environment for the combustor liner in addition to the components downstream of the combustor. Shear layers between the dilution jets and the mainstream, as well as combustor liner film-cooling interactions create a complex mean flow field within the combustor, which is not easy to model. A completely uniform temperature and velocity profile at the combustor exit is desirable from the standpoint of reducing the secondary flows in the turbine. However, this seldom occurs due to a lack of thorough mixing within the combustor. Poor mixing results in non-uniformities, such as hot streaks, and allow non-combusted fuel to exit the combustor.

This investigation developed a database documenting the thermal and flow characteristics within a combustor simulator representative of the flowfield within a gas turbine aero-engine. Three- and two-component laser Doppler velocimeter measurements were completed to quantify the flow and turbulence fields, while a thermocouple rake was used to quantify the thermal fields.

The measured results show very high turbulence levels due to the dilution flow injection. Directly downstream of the dilution jets, an increased thickness in the film-cooling was noted with a fairly non-homogeneous temperature field across the combustor width. A highly turbulent shear layer was found at the leading edge of the dilution jets. Measurements also showed that a relatively extensive recirculation region existed downstream of the dilution jets. Despite the lack of film-cooling injection at the trailing edge of the dilution hole, there existed coolant flow indicative of a horse-shoe vortex wrapping around the jet. As a result of the dilution jet interaction with the mainstream flow, kidney-shaped thermal fields and counter-rotating vortices developed. These vortices serve to enhance combustor mixing.
Master of Science

This work is concerned with the fluid dynamic processes and the associated loss of acoustic energy produced by circular apertures within noise absorbing perforated walls. Although applicable to a wide range of engineering applications particular emphasis in this work is placed on the use of such features within a gas turbine combustion system. The primary aim for noise absorbers in gas turbine combustion systems is the elimination of thermo-acoustic instabilities, which are characterised by rapidly rising pressure amplitudes which are potentially damaging to the combustion system components. By increasing the amount of acoustic energy being absorbed the occurrence of thermo-acoustic instabilities can be avoided. The fundamental acoustic characteristics relating to linear acoustic absorption are presented. It is shown that changes in orifice geometry, in terms of gas turbine combustion system representative length-to-diameter ratios, result in changes in the measured Rayleigh Conductivity. Furthermore in the linear regime the maximum possible acoustic energy absorption for a given cooling mass flow budget of a conventional combustor wall will be identified. An investigation into current Rayleigh Conductivity and aperture impedance (1D) modelling techniques are assessed and the ranges of validity for these modelling techniques will be identified. Moreover possible improvements to the modelling techniques are discussed. Within a gas turbine system absorption can also occur in the non-linear operating regime. Hence the influence of the orifice geometry upon the optimum non-linear acoustic absorption is also investigated. Furthermore the performance of non-linear acoustic absorption modelling techniques is evaluated against the conducted measurements. As the amplitudes within the combustion system increase the acoustic absorption will transition from the linear to the non-linear regime. This is important for the design of absorbers or cooling geometries for gas turbine combustion systems as the propensity for hot gas ingestion increases. Hence the relevant parameters and phenomena are investigated during the transition process from linear to non-linear acoustic absorption. The unsteady velocity field during linear and non-linear acoustic absorption is captured using particle image velocimetry. A novel analysis technique is developed which enables the identification of the unsteady flow field associated with the acoustic absorption. In this way an investigation into the relevant mechanisms within the unsteady flow fields to describe the acoustic absorption behaviour of the investigated orifice plates is conducted. This methodology will also help in the development and optimisation of future damping systems and provide validation for more sophisticated 3D numerical modelling methods. Finally a set of design tools developed during this work will be discussed which enable a comprehensive preliminary design of non-resonant and resonant acoustic absorbers with multiple perforated liners within a gas turbine combustion system. The tool set is applied to assess the impact of the gas turbine combustion system space envelope, complex swirling flow fields and the propensity to hot gas ingestion in the preliminary design stages.

Les réglementations toujours plus drastiques sur les émissions de polluants ont conduit au développement de systèmes de combustion opérant en régimes pauvres qui sont malheureusement sujet aux instabilités thermo acoustiques. La capacité de la Simulation aux Grandes Echelles (SGE) à simuler des turbines à gaz industrielles complexes de grande puissance est mise en évidence au cours de ce travail de thèse. Tout d’abord, la SGE est appliquée à un brûleur académique et validée par comparaison à des mesures effectuées à l’Université de Berlin ainsi qu’à des simulations SGE effectuées avec OpenFOAM chez Siemens. Afin de déterminer la stabilité de ce bruleur le couplage entre l’acoustique et la combustion est modélisé par l’approche de type fonction de transfert de flamme (FTF). Suite à ces calcules et l’évaluation de la FTF les fluctuations du nombre de swirl sont identifiées comme un paramètre à même de modifier cette réponse de flamme. Après cette première étape de validation, une turbine à gaz industrielle est simulée en SGE pour deux géométries différentes du brûleur et pour deux points de fonctionnement. La FTF issue de ces calculs est peu influencée par les deux points de fonctionnement. A l’inverse, des légères modifications de la géométrie du swirler modifient les caractéristiques de la FTF montrant que plusieurs mécanismes sont en jeu. Ces mécanismes sont identifiés comme étant la vitesse d’entrée, les fluctuations de swirl et les fluctuations de fraction de mélange. Cette dernière est causée par: 1) la pulsation du débit de carburant injecté et 2) la trajectoire fluctuante des jets de carburant. Bien que le swirler soit conçu pour fournir un mélange le plus homogène possible, d’importantes hétérogénéités de mélange à l’entrée de la chambre de combustion sont présentes. Les perturbations de mélange se combinent avec les fluctuations de vitesse (et donc avec les fluctuations de swirl) aboutissant à des résultats de FTF différents. Un modèle étendu pour la FTF reliant le dégagement de chaleur à la vitesse d’entrée et à la fluctuation de fraction de mélange (modèle MISO) se révèle être une bonne solution pour ces systèmes complexes. Une analyse non linéaire montre en outre que l’amplitude de forçage conduit non seulement à une saturation de la flamme, mais aussi à un changement de la réponse de flamme. La saturation de la flamme n’est vérifiée que pour la FTF globale et le gain augmente localement avec une amplitude croissante. Pour ce système on notera enfin que la flamme linéaire, comme la flamme non linéaire, ne sont pas compactes: certaines zones pourtant situées l’une à coté de l’autre, ont des différences significatives de délai de FTF, montrant que certaines parties de la flamme amortissent l’excitation alors que d’autres l’amplifient
Modern pollutant regulation have led to a trend towards lean combustion systems which are prone to thermo-acoustic instabilities. The ability of Large Eddy Simulation (LES) to handle complex industrial heavy-duty gas turbines is evidenced during this thesis work. First, LES is applied to an academic single burner in order to validate the modeling against measurements performed at TU Berlin and against OpenFoam LES simulations done at Siemens. The coupling between acoustic and combustion is modeled with the Flame Transfer Function (FTF) approach and swirl number fluctuations are identified changing the FTF amplitude response of the flame. Then, an industrial gas turbine is analyzed for two different burner geometries and operating conditions. The FTF is only slightly influenced for the two operating points but slight modifications of the swirler geometry do modify the characteristics of the FTF showing that a simple model taking only into account the flight time is not appropriate and additional mechanisms are at play. Those mechanisms are identified being the inlet velocity, the swirl and the inlet mixture fraction fluctuations. The latter is caused by two mechanisms: 1) the pulsating injected fuel flow rate and 2) the fluctuating trajectory of the fuel jets. Although the diagonal swirler is designed to provide good mixing, effects of mixing heterogeneities at the combustion chamber inlet occur. Mixture perturbations phase with velocity (and hence with swirl) fluctuations and combine with them to lead to different FTF results. Another FTF approach linking heat release to inlet velocity and mixture fraction fluctuation (MISO model) shows further to be a good solution for complex systems. A nonlinear analysis shows that the forcing amplitude not only leads to a saturation of the flame but also to changes of the delay response. Flame saturation is only true for the global FTF and the gain increases locally with increasing forcing amplitude. Both, the linear and the nonlinear flames, are not compact: flame regions located right next to each other exhibited significant differences in delay meaning that at the same instant certain parts of the flame damp the excitation while others feed it

The world demand for electrical power increasescontinuously, requiring efficient and low-cost methods forpower generation. This thesis investigates two advanced powercycles with mixtures as the working fluid: the Kalina cycle,alternatively called the ammonia-water cycle, and theevaporative gas turbine cycle. These cycles have the potentialof improved performance regarding electrical efficiency,specific power output, specific investment cost and cost ofelectricity compared with the conventional technology, sincethe mixture working fluids enable efficient energyrecovery.

This thesis shows that the ammonia-water cycle has a betterthermodynamic performance than the steam Rankine cycle as abottoming process for natural gas-fired gas and gas-dieselengines, since the majority of the ammonia-water cycleconfigurations investigated generated more power than steamcycles. The best ammonia-water cycle produced approximately40-50 % more power than a single-pressure steam cycle and 20-24% more power than a dual-pressure steam cycle. The investmentcost for an ammonia-water bottoming cycle is probably higherthan for a steam cycle; however, the specific investment costmay be lower due to the higher power output.

A comparison between combined cycles with ammonia-waterbottoming processes and evaporative gas turbine cycles showedthat the ammonia-water cycle could recover the exhaust gasenergy of a high pressure ratio gas turbine more efficientlythan a part-flow evaporative gas turbine cycle. For a mediumpressure ratio gas turbine, the situation was the opposite,except when a complex ammonia-water cycle configuration withreheat was used. An exergy analysis showed that evaporativecycles with part-flow humidification could recover energy asefficiently as, or more efficiently than, full-flow cycles. Aneconomic analysis confirmed that the specific investment costfor part-flow cycles was lower than for full-flow cycles, sincepart-flow humidification reduces the heat exchanger area andhumidification tower volume. In addition, the part-flow cycleshad lower or similar costs of electricity compared with thefull-flow cycles. Compared with combined cycles, the part-flowevaporative cycles had significantly lower total and specificinvestment costs and lower or almost equal costs ofelectricity; thus, part-flow evaporative cycles could competewith the combined cycle for mid-size power generation.

Keywords:power cycle, mixture working fluid, Kalinacycle, ammonia-water mixture, reciprocating internal combustionengine, bottoming cycle, gas turbine, evaporative gas turbine,air-water mixture, exergy

Au cours des dernières décennies, les instabilités de combustion ont constitué un important défi pour de nombreux projets industriels, en particulier dans la conception de moteurs-fusées à ergols liquide et de turbines à gaz. L'atténuation de leurs effets nécessite une solide compréhension scientifique de l'interaction complexe entre la dynamique de flamme et les ondes acoustiques qu'elles impliquent. Au cours de cette thèse, plusieurs directions ont été explorées pour fournir une meilleure compréhension de la dynamique des flammes dans les moteurs-fusées cryogéniques, ainsi que des méthodes numériques plus efficaces et robustes pour la prédiction des instabilités thermoacoustiques dans les chambres de combustion à géométries complexes. La première facette de ce travail a consisté en la résolution de modes thermoacoustiques dans les chambres de combustion complexes comportant à injecteurs multiples, une tâche qui nécessite souvent des simplifications pour être abordable en termes de coût de calcul. Ces hypothèses physiques nécessaires ont conduit à la popularité croissante des modèles bas-ordre acoustiques, parmi lesquels ceux utilisant l'expansion de Galerkin ont démontré une efficacité prometteuse tout en conservant une précision satisfaisante. Ceux-ci sont cependant limités à des géométries simples qui n'intègrent pas les caractéristiques complexes des systèmes industriels. Une grande partie de ce travail a donc consisté tout d'abord à identifier clairement les limitations mathématiques de l'expansion classique de Galerkin, puis à concevoir un nouveau type d'expansion modale, appelé expansion sur frame, qui ne souffre pas des mêmes restrictions. En particulier, l'expansion sur frame est capable de représenter avec précision le champ de vitesse acoustique près des parois de la chambre de combustion autres que des murs rigides, une capacité cruciale qui manque à la méthode Galerkin. Dans ce travail, le concept d'expansion modale de surface a également été introduit pour modéliser des frontières topologiquement complexes, comme les plaques multi-perforées rencontrées dans les turbines à gaz. Ces nouvelles méthodes numériques ont été combinées avec le formalisme state-space pour construire des réseaux acoustiques de systèmes complexes. Le modèle obtenu a été implémenté dans le code STORM (State-space Thermoacoustic low-ORder Model), qui permet la modélisation bas-ordre des instabilités thermoacoustiques dans des géométries arbitrairement complexes. Le deuxième ingrédient de la prédiction des instabilités thermoacoustiques est la modélisation de la dynamique de flamme. Ce travail a traité de ce point, dans le cas spécifique d'une flamme-jet cryogénique caractéristique d'un moteur-fusée à ergols liquides. Les phénomènes contrôlant la dynamique de flamme ont été identifiés grâce à des Simulations aux Grandes Échelles (SGE) du banc d'essai expérimental Mascotte, où les deux réactifs (CH4 et O2) sont injectés dans des conditions transcritiques. Une première simulation donne un aperçu détaillé de la dynamique intrinsèque de la flamme. Plusieurs SGE avec modulation harmonique de l'injection de carburant, à différentes fréquences et amplitudes, ont été effectués afin d'évaluer la réponse de la flamme aux oscillations acoustiques et de calculer une Fonction de Transfert de Flamme (FTF). La réponse non-linéaire de la flamme, notamment les interactions entre les oscillations intrinsèques et forcées, a également été étudiée. Enfin, la stabilisation de cette flamme dans la région proche de l'injecteur, qui est d'une importance primordiale sur la dynamique globale de la flamme, a été étudiée grâce à une simulation directe multi-physique, où un problème couplé de transfert de chaleur est résolu au niveau de la lèvre de l'injecteur
Over the last decades, combustion instabilities have been a major concern for a number of industrial projects, especially in the design of Liquid Rocket Engines (LREs) and gas turbines. Mitigating their effects requires a solid scientific understanding of the intricate interplay between flame dynamics and acoustic waves that they involve. During this PhD work, several directions were explored to provide a better comprehension of flame dynamics in cryogenic rocket engines, as well as more efficient and robust numerical methods for the prediction of thermoacoustic instabilities in complex combustors. The first facet of this work consisted in the resolution of unstable thermoacoustic modes in complex multi-injectors combustors, a task that often requires a number of simplifications to be computationally affordable. These necessary physics-based assumptions led to the growing popularity of acoustic Low-Order Models (LOMs), among which Galerkin expansion LOMs have displayed a promising efficiency while retaining a satisfactory accuracy. Those are however limited to simple geometries that do not incorporate the complex features of industrial systems. A major part of this work therefore consisted first in clearly identifying the mathematical limitations of the classical Galerkin expansion, and then in designing a novel type of modal expansion, named a frame expansion, that does not suffer from the same restrictions. In particular, the frame expansion is able to accurately represent the acoustic velocity field, near non-rigid-wall boundaries of the combustor, a crucial ability that the Galerkin method lacks. In this work, the concept of surface modal expansion is also introduced to model topologically complex boundaries, such as multi-perforated liners encountered in gas turbines. These novel numerical methods were combined with the state-space formalism to build acoustic networks of complex systems. The resulting LOM framework was implemented in the code STORM (State-space Thermoacoustic low-ORder Model), which enables the low-order modeling of thermoacoustic instabilities in arbitrarily complex geometries. The second ingredient in the prediction of thermoacoustic instabilities is the flame dynamics modeling. This work dealt with this problem, in the specific case of a cryogenic coaxial jet-flame characteristic of a LRE. Flame dynamics driving phenomena were identified thanks to three-dimensional Large Eddy Simulations (LES) of the Mascotte experimental test rig where both reactants (CH4 and O2) are injected in transcritical conditions. A first simulation provides a detailed insight into the flame intrinsic dynamics. Several LES with harmonic modulation of the fuel inflow at various frequencies and amplitudes were performed in order to evaluate the flame response to acoustic oscillations and compute a Flame Transfer Function (FTF). The flame nonlinear response, including interactions between intrinsic and forced oscillations, were also investigated. Finally, the stabilization of this flame in the near-injector region, which is of primary importance on the overall flame dynamics, was investigated thanks to muulti-physics two-dimensional Direct Numerical Simulations (DNS), where a conjugate heat transfer problem is resolved at the injector lip

Competitive pressure and stringent emissions legislation have placed an urgent demand on research to improve our understanding of the gas turbine combustor flow field. Flow through the air admission ports of a combustor plays an essential role in determining the internal flow patterns on which many features of combustor performance depend. This thesis explains how a combination of experimental and computational research has helped improve our understanding, and ability to predict, the flow characteristics of jets entering a combustor. The experiments focused on a simplified generic geometry of a combustor port system. Two concentric tubes, with ports introduced into the inner tube's wall, allowed a set of radially impinging jets to be formed within the inner tube. By investigating the flow with LDA instrumentation and flow visualisation methods a quantitative and qualitative picture of the mean and turbulent flow fields has been constructed. Data were collected from the annulus, port and core regions. These data provide suitable validation information for computational models, allow improved understanding of the detailed flow physics and provide the global performance parameters used traditionally by combustor designers. Computational work focused on improving the port representation within CFD models. This work looked at the effect of increasing the grid refinement, and improving the geometrical representation of the port. The desire to model realistic port features led to the development of a stand-alone port modelling module. Comparing calculations of plain-circular ports to those for more realistic chuted port geometry, for example, showed that isothermal modelling methods were able to predict the expected changes to the global parameters measured. Moreover, these effects are seen to have significant consequences on the predicted combustor core flow field.

Gas turbine combustor design represents an ambitious task in numerical and experimental analysis. A significant number of competing criteria must be optimised within specified constraints in order to satisfy legislative and performance requirements. Currently, preliminary combustor flow and heat transfer design procedures, which by necessity involve semi-empirical models, are often restricted in their range of application. The objective of this work is the development of a versatile design tool able to model all conceivable gas turbine combustor types. A network approach provides the foundation for a complete flow and heat transfer analysis to meet this goal. The network method divides the combustor into a number of independent interconnected sub-flows. A pressure-correction methodology solves the continuity equation and a pressure-drop/flow-rate relationship. A constrained equilibrium calculation, incorporating mixing and recirculation models, simulates the combustion process. The new procedures are validated against numerical and experimental data within three annular combustors and one reverse flow combustor. A full conjugate heat transfer model is developed to allow the calculation of liner wall temperature characteristics. The effects of conduction, convection and radiation are included in the model. Film cooling and liner heat pick-up effects are included in the convection calculation. Radiation represents the most difficult mode of heat transfer to simulate in the combustion environment. A discrete transfer radiation model is developed and validated for use within the network solver. The effects of soot concentration on radiation is evaluated with the introduction of radial properties profiles. The accuracy of the heat transfer models are evaluated with comparisons to experimental thermal paint temperature data on a reverse flow and annular combustors. The resulting network analysis code represents a powerful design tool for the combustion engineer incoporating a novel and unique strategy.

Finding limit cycles and their stability is one of the central problems of nonlinear thermoacoustics. However, a limit cycle is not the only type of self-excited oscillation in a nonlinear system. Nonlinear systems can have quasi-periodic and chaotic oscillations. This thesis examines the different types of oscillation in a numerical model of a ducted premixed flame, the bifurcations that lead to these oscillations and the influence of external forcing on these oscillations. Criteria for the existence and stability of limit cycles in single mode thermoacoustic systems are derived analytically. These criteria, along with the flame describing function, are used to find the types of bifurcation and minimum triggering amplitudes. The choice of model for the velocity perturbation field around the flame is shown to have a strong influence on the types of bifurcation in the system. Therefore, a reduced order model of the velocity perturbation field in a forced laminar premixed flame is obtained from Direct Numerical Simulation. It is shown that the model currently used in the literature precludes subcritical bifurcations and multi-stability. The self-excited thermoacoustic system is simulated in the time domain with many modes in the acoustics and analysed using methods from nonlinear dynamical systems theory. The transitions to the periodic, quasiperiodic and chaotic oscillations are via sub/supercritical Hopf, Neimark-Sacker and period-doubling bifurcations. Routes to chaos are established in this system. It is shown that the single mode system, which gives the same results as a describing function approach, fails to capture the period-$2$, period-$k$, quasi-periodic and chaotic oscillations or the bifurcations and multi-stability seen in the multi-modal case, and underpredicts the amplitude. Instantaneous flame images reveal that the wrinkles on the flame surface and pinch off of flame pockets are regular for periodic oscillations, while they are irregular and have multiple time and length scales for quasi-periodic and chaotic oscillations. Cusp formation, their destruction by flame propagation normal to itself, and pinch-off and rapid burning of pockets of reactants are shown to be responsible for generating a heat release rate that is a highly nonlinear function of the velocity perturbations. It is also shown that for a given acoustic model of the duct, many discretization modes are required to capture the rich dynamics and nonlinear feedback between heat release and acoustics seen in experiments. The influence of external harmonic forcing on self-excited periodic, quasi-periodic and chaotic oscillations are examined. The transition to lock-in, the forcing amplitude required for lock-in and the system response at lock-in are characterized. At certain frequencies, even low-amplitude forcing is sufficient to suppress period-$1$ oscillations to amplitudes that are 90$\%$ lower than that of the unforced state. Therefore, open-loop forcing can be an effective strategy for the suppression of thermoacoustic oscillations. This thesis shows that a ducted premixed flame behaves similarly to low-dimensional chaotic systems and that methods from nonlinear dynamical systems theory are superior to the describing function approach in the frequency domain and time domain analysis currently used in nonlinear thermoacoustics.

Thesis (M.S. in Aeronautical Engineering)--Naval Postgraduate School, March 1990.
Thesis Advisor(s): Netzer, D. W. Second Reader: Shreeve, R. P. "March 1990." Description based on signature page as viewed on August 25, 2009. DTIC Descriptor(s): Air, augmentation, catalysis, catalysts, combustors, configurations, engines, fuels, gas generating systems, gas turbines, measurement, pressure, profiles, ratios, reduction, tubes, velocity. DTIC Identifier(s): Nitrogen oxides, NOx control, Gas turbines, Gas analyzers, Pollution abatement, computer programs, Emissions control, Exhaust augmentor tubes, Thesis. Author(s) subject terms: Nox control, gas turbine combustors; emissions control exhaust augmentor tubes; gas analyzers; pollution control. Includes bibliographical references (p. 73-74). Also available online.

Дисертація на здобуття наукового ступеня кандидата технічних наук за спе-ціальністю 05.05.03 – двигуни та енергетичні установки. – Національний технічний університет «Харківський політехнічний інститут». – Харків, 2017. Дисертаційна робота присвячена вибору схеми та параметрів силової енергетичної установки для утилізації попутного нафтового газу. В роботі проаналізовано можливість використання різних силових установок для утилізації попутного нафтового газу. Розроблено схеми енергогенеруючих установок з використанням газо-турбінних і газопоршневих двигунів внутрішнього згоряння для виробництва еле-ктричної енергії за рахунок утилізації попутного нафтового газу на нафтовидобувних і нафтопереробних підприємств. Використано енергоексергетичний метод для оцінки ефективності запропонованих схем. Проведено економічний аналіз доцільності побудови енергогенеруючих потужностей, що будуть споживати попутний нафтовий газ з аналізом чутливості для таких параметрів як зміна ціни на електроенергію та вплив високих температур навколишнього середовища. При зміні температури навколишнього середовища з +15 до +45 °С кількість енергії, що виробляється для проекту А буде зменшуватися на 26%, для проекту В – на 10,9%. Визначено, що попри більшу вартість проекту В ($2 843 009.55) супроти проекту А ($1 964 434.69), термін окупності складає: для проекту А – 6 років, 1 місяць; для проекту В – 3 роки, 8 місяців. Обґрунтована доцільність використання поршневого ДВЗ у складі енергогенеруючої установки.
Thesis for the degree of candidate of technical sciences by specialty 05.05.03 - engines and power units. - National Technical University "Kharkiv Polytechnic Institute". - Kharkiv, 2017. The thesis is devoted to the choice of the scheme and parameters of the power plant for utilization of associated petroleum gas. The paper analyzes the possibility of using various power plants for utilization of associated petroleum gas. The schemes of power units using gas turbine and gas piston internal combustion engines to generate power electricity have been developed by using associated petroleum gas in oil refinery. The anergy-exergy method was used to analys the effectiveness of the proposed schemes. An economic analysis of the feasibility of constructing power generating capacities that will consume the associated oil gas with an analysis of sensitivity for such parameters as a change in the price of electricity and the impact of high ambient temperatures has been carried out. If the ambient temperature is changed from +15 to + 45 ° C, the amount of energy generated for Project A will be reduced by 26%, for Project B - by 10.9%. It is determined that despite the high cost of Project B ($ 2,843,009.55) against Project A ($ 1964,434.69), the payback period is: for Project A - 6 years, 1 month; For Project B - 3 years, 8 months. The expediency of using the piston-internal ICE as part of the power generating unit is substantiated.

Stationary gas turbines manufacturers and operators are under constant scrutiny to both reduce environmentally harmful emissions and obtain efficient combustion. Numerical simulations have become an integral part of the development and optimisation of gas turbine combustors. In this thesis work, the gas turbine combustion process is analysed in two parts, a study on air-fuel mixing and turbulent combustion. For computational fluid dynamic analysis work the open-source CFD code OpenFOAM and STAR-CCM+ are used. A fuel jet injected to cross-flowing air flow is simplified air-fuel mixing arrangement, and this problem is analysed numerically in the first part of the thesis using both Reynolds Averaged Navier Stokes (RANS) method and Large Eddy Simulation (LES) methods. Several turbulence models are compared against experimental data in this work, and the complex turbulent vortex structures their effect on mixing field prediction is observed. Furthermore, the numerical methods are extended to study twin jets in cross-flow interaction which is relevant in predicting air-fuel mixing with arrays of fuel injection nozzles. LES methods showed good results by resolving the complex turbulent structures, and the interaction of two jets is also visualised. In this work, all three turbulent combustion regimes non-premixed, premixed, partially premixed are modelled using different combustion models. Hydrogen blended fuels have drawn particular interest recently due to enhanced flame stabilisation, reduced CO2 emissions, and is an alternative method to store energy from renewable energy sources. Therefore, the well known Sydney swirl flame which uses CH4: H2 blended fuel mixture is modelled using the steady laminar flamelet model. This flame has been found challenging to model numerically by previous researchers, and in this work, this problem has been addressed with improved combustion modelling approach with tabulated chemistry. Recognizing that the current and future gas turbine combustors operate on a mixed combustion regime during its full operational cycle, combustion simulations of premixed/partially premixed flames are also performed in this thesis work. Dynamical artificially thickened flame model is implemented in OpenFOAM and validated using propagating and stationary premixed flames. Flamelet Generated Manifold (FGM) methods are used in the modelling of turbulent stratified flames which is a relatively new field of under investigation, and both experimental and numerical analysis is required to understand the physics. The recent experiments of the Cambridge stratified burner are studied using the FGM method in this thesis work, and good agreement is obtained for mixing field and temperature field predictions.

Substantial economic and even safety related gains can be achieved if effective gas turbine performance analysis is attained. During the development phase, analysis can help understand the effect on the various components and on the overall engine performance of the modifications applied. During usage, analysis plays a major role in the assessment of the health status of the engine. Both condition monitoring of operating engines and pass off tests heavily rely on the analysis. In spite of its relevance, accurate performance analysis is still difficult to achieve. A major cause of this is measurement uncertainty: gas turbine measurements are affected by noise and biases. The simultaneous presence of engine and sensor faults makes it hard to establish the actual condition of the engine components. To date, most estimation techniques used to cope with measurement uncertainty are based on Kalman filtering. This classic estimation technique, though, is definitely not effective enough. Typical Kalman filter results can be strongly misleading so that even the application of performance analysis may become questionable. The main engine manufactures, in conjunction with research teams, have devised modified Kalman filter based techniques to overcome the most common drawbacks. Nonetheless, the proposed methods are not able to produce accurate and reliable performance analysis. In the present work a different approach has been pursued and a novel method developed, which is able to quantify the performance parameter variations expressing the component faults in presence of noise and a significant number of sensor faults. The statistical basis of the method is sound: the only accepted statistical assumption regards the well known measurement noise standard deviations. The technique is based on an optimisation procedure carried out by means of a problem specific, real coded Genetic Algorithm. The optimisation based method enables to concentrate the steady state analysis on the faulty engine component(s). A clear indication is given as to which component(s) is(are) responsible for the loss of performance. The optimisation automatically carries out multiple sensor failure detection, isolation and accommodation. The noise and biases affecting the parameters setting the operating point of the engine are coped with as well. The technique has been explicitly developed for development engine test bed analysis, where the instrumentation set is usually rather comprehensive. In other diagnostic cases (pass off tests, ground based analysis of on wing engines), though, just few sensors may be present. For these situations, the standard method has been modified to perform multiple operating point analysis, whereby the amount of information is maximised by simultaneous analysis of more than a single test point. Even in this case, the results are very accurate. In the quest for techniques able to cope with measurement uncertainty, Neural Networks have been considered as well. A novel Auto-Associative Neural Network has been devised, which is able to carry out accurate sensor failure detection and isolation. Advantages and disadvantages of Neural Network-based gas turbine diagnostics have been analysed.