No description available.

No description available.

No description available.

Trapped Vortex Combustor (TVC) is a relatively new concept for potential use in gas turbine engines addressing ever increasing demands of high efficiency, low emissions, low pressure drop, and improved pattern factor. This concept holds promise for future because of its inherent advantages over conventional swirl-stabilized combustors. The main difference between TVC and a conventional gas turbine combustor is in the way combustion is stabilized. In conventional combustors, flame is stabilized because of formation of toroidal flow pattern in the primary zone due to interaction between incoming swirling air and fuel flow. On the other hand, in TVC, there is a physical cavity in the wall of combustor with continuous injection of air and fuel leading to stable and sustained combustion. Past work related to TVC has focussed on use of two cavities in the combustor liner. In the present study, a single cavity combustor concept is evaluated through simulation and experiments for applications requiring compact combustors such as Unmanned Aerial Vehicles (UAVs) and cruise missiles. In the present work, numerical simulations were initially performed on a planar, rectangular single-cavity geometry to assess sensitivity of various parameters and to design a single-cavity TVC test rig. A water-cooled, modular, atmospheric pressure TVC test rig is designed and fabricated for reacting and non-reacting flow experiments. The unique features of this rig consist of a continuously variable length-to-depth ratio (L/D) of the cavity and optical access through quartz plates provided on three sides for visualization. Flame stabilization in the single cavity TVC was successfully achieved with methane as fuel, and the range of flow conditions for stable operation were identified. From these, a few cases were selected for detailed experimentation. Reacting flow experiments for the selected cases indicated that reducing L/D ratio and increasing cavity-air velocity favour stable combustion. The pressure drop across the single-cavity TVC is observed to be lower as compared to conventional combustors. Temperatures are measured at the exit using thermocouples and corrected for radiative losses. Species concentrations are measured at the exit using an exhaust gas analyzer. The combustion efficiency is observed to be around 98-99% and the pattern factor is observed to be in the range of 0.08 to 0.13. High-speed imaging made possible by the optical access indicates that the overall combustion is fairly steady, and there is no major vortex shedding downstream. This enabled steady-state simulations to be performed for the selected cases. Insight from simulations has highlighted the importance of air and fuel injection strategies in the cavity. From a mixing and combustion efficiency standpoint, it is desirable to have a cavity vortex that is anti-clockwise. However, the natural tendency for flow over a cavity is to form a vortex that is clockwise. The tendency to blow-out at higher inlet flow velocities is thought to be because of these two opposing effects. This interaction helps improve mixing, however leads to poor flame stability unless cavity-air velocity is strong enough to support a strong anti-clockwise vortex in the cavity. This basic understating of cavity flow dynamics can be used for further design improvements in future to improve flame stability at higher inlet flow velocities and eventually lead to the development of a practical combustor.

In many areas of aerodynamics the technique of Large Eddy Simulation (LES) has proved a practical way of modelling the unsteady phenomena in numerical simulations. Few applications are as dependent on such an approach as the prediction of flow within a gas turbine combustor. Like any form of Computational Fluid Dynamics (CFD), LES requires specification of the velocity field at the inflow boundary, with much evidence suggesting the specification of inlet turbulence can be critical to the resultant accuracy of the prediction. While a database of time-resolved velocity data may be obtained from a precursor LES calculation, this technique is prohibitively expensive for complex geometries. An alternative is to use synthetic inlet conditions obtained from experimental data High-speed Particle Image Velocimetry (PIV) is used here to provide planar velocity data at up to 1kHz temporal resolution in two test cases representative of gas turbine combustor flows (a vortex generator in a duct and an idealised combustor). As the data sampling rate is approaching a typical LES time-step it introduces the possibility of applying instantaneous experimental data directly as an inlet condition. However, as typical solution domain inlet regions for gas turbine combustor geometries cannot be adequately captured in a single field of PIV data, it is necessary to consider a method by which a synchronous velocity field may be obtained from multiple PIV fields that were not captured concurrently. A method is proposed that attempts to achieve this by a combined process of Linear Stochastic Estimation and high-pass filtering. The method developed can be generally applied without a priori assumptions of the flow and is demonstrated to produce a velocity field that matches very closely that of the original PIV, with no discontinuities in the velocity correlations. The fidelity and computational cost of the method compares favourably to several existing inlet condition generation methods. Finally, the proposed and existing methods for synthetic inlet condition generation are applied to LES predictions of the two test cases. There is shown to be significant differences in the resulting flow, with the proposed method showing a marked ii reduction in the adjustment period that is required to establish turbulent equilibrium downstream of the inlet. However, it is noted the presence of downstream turbulence generating features can mask any differences in the inlet condition, to the extent that the flow in the core of the combustor test case is found to be insensitive to the inlet condition applied at the entry to the feed annulus for the test conditions applied here.

It has been recognized that the evaporation process is one of the pivotal mechanisms driving thermo-acoustic instability in gas turbines and rockets in particular. In this regard, this study is principally focused on studying the evaporation process relevant to thermo-acoustic instability from three complementary viewpoints in an effort to contribute to an overall instability model driven primarily by evaporation in gas turbine combustors. Firstly, a state of the art LES algorithm is employed to validate an evaporation model to be employed in predictive modelling regarding combustion instabilities. Good agreement between the numerical predictions and experimental data is achieved. Additionally, transient sub-critical droplet evaporation is investigated numerically. In particular, a numerical method is proposed to capture the extremely important pressure-velocity-density coupling. Furthermore, the dynamic system nonlinear behaviour encountered in classical thermo-acoustic instability is investigated. The Poincaré map is adopted to analyse the stability of a simple non-autonomous system considering a harmonic oscillation behaviour for the combustion environment. The bifurcation diagram of a one-mode model is obtained where the analysis reveals a variety of chaotic behaviours for some select ranges of the bifurcation parameter. The bifurcation parameter and the corresponding period of a two-mode dynamic model are calculated using both analytical and numerical methods. The results computed by different methods are in good agreement. In addition, the dependence of the bifurcation parameter and the period on all the relevant coefficients in the model is investigated in depth. Moreover, a discrete dynamic model accounting for both combustion and vaporization processes is developed. In terms of different bifurcation parameters relevant to either combustion or evaporation, various bifurcation diagrams are presented. As part of the nonlinear characterization, the governing process Lyapunov exponent is calculated and employed to analyze the stability of the particular dynamic system. The study has shown conclusively that the evaporation process has a significant impact on the intensity and nonlinear behaviour of the system of interest, vis-à-vis a model accounting for only the gaseous combustion process. Furthermore, two particular nonlinear control methodologies are adopted to control the chaotic behaviour displayed by the particular aperiodic motions observed. These algorithms are intended to be implemented for control of combustion instability numerically and experimentally to provide a rational basis for some of the control methodologies employed in the literature. Finally, a state of the art neural network is employed to identify and predict the nonlinear behaviour inherent in combustion instability, and control the ensuing pressure oscillations. Essentially, the NARMAX model is implemented to capture nonlinear dynamics relating the input and output of the system of interest. The simulated results accord with the results reported. Moreover, a control system using the NARMA-L2 algorithm is developed. The simulation conclusively points to the fact that the amplitude of pressure oscillations can be attenuated to an acceptable level and the controller proposed may be implemented in a practical manner.

No description available.

The complex multiphase flow physics of spray-swirl interaction in both reacting and non-reacting environment is of fundamental and applied significance for a wide variety of applications ranging from gas turbine combustors to pharmaceutical drug nebulizers. Understanding the intricate dynamics between this two phase flow field is pivotal for enhancing mixing characteristics, reducing pollutant emissions and increasing the combustion efficiency in next generation combustors. The present work experimentally investigates the near and far-field break-up, dispersion and coalescence characteristics of a hollow cone spray in an unconfined, co¬annular isothermal swirling air jet environment. The experiments were conducted using an axial-flow hollow cone spray nozzle having a 0.5 mm orifice. Nozzle injection pressure (PN = 1 bar) corresponding to a Reynolds number at nozzle exit ReN = 7900 used as the test setting. At this setting, the operating Reynolds number of the co-annular swirling air stream number (Res) was varied in four distinct steps, i.e. Res = 1600, 3200, 4800 and 5600. Swirl was imparted to the co¬axial flow using a guided vane swirler with blade angle of Ф=45° (corresponding geometric swirl number SG = 0.8). Two types of laser diagnostic techniques were utilized: Particle / Droplet imaging analysis (PDIA) and shadowgraph to study the underlying physical mechanisms involved in the primary breakup, dispersion and coalescence dynamics of the spray. Measurements were made in the spray in both axial and radial directions and they indicate that Sauter Mean Diameter (SMD) in radial direction is highly reliant on the intensity of swirl imparted to the spray. The spray is subdivided into two zones as function of swirl in axial and radial direction: (1) near field of the nozzle (ligament regime) where variation in SMD arises predominantly due to primary breakup of liquid films (2) far-field of the nozzle where dispersion and collision induced coalescence of droplets is dominant. Each regime has been analyzed meticulously, by computing probability of primary break-up, probability of coalescence and spatio-temporal distribution of droplets which gives probabilistic estimate of aforementioned governing phenomena. In addition to this, spray global length scale parameters such as spray cone angle, break-up length, wavelength of liquid film has been characterized by varying Res while maintaining constant ReN.

<div>The aviation industry is continuously growing amid tight restrictions on global emission</div><div>reductions. Alternative aviation fuels have gained attention and developed to replace the</div><div>conventional petroleum-derived aviation fuels. The replacement of conventional fuels with</div><div>alternative fuels, which are composed solely of hydrocarbons (non-petroleum), can mitigate</div><div>impacts on the environment and diversify the energy supply, potentially reducing fuel costs.</div><div>To ensure the performance of alternative fuels, extensive laboratory and full-scale engine</div><div>testings are required, thereby a lengthy and expensive process. The National Jet Fuel Combustion</div><div>Program (NJFCP) proposed a plan to reduce this certification process time and</div><div>the cost dramatically by implementing a computational model in the process, which can be</div><div>replaced with some of the testings. This requires an understanding of the influence of chemical/</div><div>physical properties of alternative fuels on combustion performance. The main objective</div><div>of this work is to investigate the spray characteristics of alternative aviation fuels compared</div><div>to that of conventional aviation fuels, which have been characterized by different physical</div><div>liquid properties at different gas turbine-relevant conditions.</div><div>The experimental work focuses on the spray characteristics of standard and alternative</div><div>aviation fuels at three operating conditions such as near lean blowout (LBO), cold engine</div><div>start, and high ambient pressure conditions. The spray generated by a hybrid pressureswirl</div><div>airblast atomizer was investigated by measuring the drop size and drop velocity at</div><div>a different axial distance downstream of the injector using a phase Doppler anemometry</div><div>(PDA) measurement system. This provided an approximate trajectory of the largest droplet</div><div>as it traveled down from the injector. At LBO conditions, the trend of decreasing drop size</div><div>and increasing drop velocity with an increase in gas pressure drop was observed for both</div><div>conventional (A-2) and alternative aviation fuels (C-1, C-5, C-7, and C-8), while the effect of</div><div>fuel injection pressure on the mean drop size and drop velocity was observed to be limited.</div><div>Moreover, the high-speed shadowgraph images were also taken to investigate the effect of</div><div>the pressure drop and fuel injection pressures on the cone angles. Their effects were found</div><div>to be limited on the cone angle.</div><div><div>The spray characteristics of standard (A-2 and A-3) and alternative (C-3) fuels were</div><div>investigated at engine cold-start conditions. At such a crucial condition, sufficient atomization</div><div>needs to be maintained to operate the engine properly. The effect of fuel properties,</div><div>especially the viscosity, was investigated on spray drop size and drop velocity using both</div><div>conventional and alternative aviation fuels. The effect of fuel viscosity was found to be minimal</div><div>and dominated by the effect of the surface tension, even though it showed a weak trend</div><div>of increasing drop size with increasing surface tension. The higher swirler pressure drop</div><div>reduced the drop size and increased drop velocity due to greater inertial force of the gas for</div><div>both conventional and alternative aviation fuels at the cold start condition. However, the</div><div>effect of pressure drop was observed to be reduced at cold start condition compared to the</div><div>results from the LBO condition.</div><div>The final aspect of experimental work focuses on the effect of ambient pressures on the</div><div>spray characteristics for both conventional (A-2) and alternative (C-5) aviation fuels. Advanced</div><div>aviation technology, especially in turbomachinery, has resulted in a greater pressure</div><div>ratio in the compressor; therefore, greater pressure in combustors for better thermal efficiency.</div><div>The effect of ambient pressure on drop size, drop velocity, and spray cone angle was</div><div>investigated using the PDA system and simultaneous Planar Laser-Induced Fluorescence</div><div>(PLIF) and Mie scattering measurement. A significant reduction in mean drop size was</div><div>observed with increasing ambient pressure, up to 5 bar. However, the reduction in the mean</div><div>drop size was found to be limited with a further increase in the ambient pressure. The effect</div><div>of the pressure drop across the swirler was observed to be significant at ambient pressure of</div><div>5 bar. The spray cone angle estimation at near the swirler exit and at 25.4 mm downstream</div><div>from the swirler exit plane using instantaneous Mie images was found to be independent of</div><div>ambient pressure. However, the cone angle at measurement plane of 18 mm in the spray</div><div>was observed to increase with increasing ambient pressure due to entrainment of smaller</div><div>droplets at higher ambient pressure. Furthermore, the fuel droplet and vapor distribution in</div><div>the spray were imaged and identified by comparing instantaneous PLIF and Mie images.</div><div>Lastly, a semi-empirical model was also developed using a phenomenological three-step</div><div>approach for the atomization process of the hybrid pressure-swirl airblast atomizer. This</div><div>model includes three sub-models: pressure-swirl spray droplet formation, droplet impingement, and film formation, and aerodynamic breakup. The model predicted drop sizes as a</div><div>function of ALR, atomizing gas velocity, surface tension, density, and ligament length and</div><div>diameter and successfully demonstrated the drop size trend observed with fuel viscosity,</div><div>surface tension, pressure drop, and ambient pressure. The model provided insights into the</div><div>effect of fuel properties and engine operating parameters on the drop size. More experimental</div><div>work is required to validate the model over a wider range of operating conditions and</div><div>physical fuel properties.</div><div>Overall, this work provides valuable information to increase understanding of the spray</div><div>characteristics of conventional and alternative aviation fuels at various engine operating</div><div>conditions. This work can provide valuable data for developing an advanced computational</div><div>combustor model, ultimately expediting the certification of new alternative aviation fuels.</div></div>

The modern gas turbine engine industry needs a simpler and faster method to facilitate the design of gas turbine combustors due to the enormous costs of experimental test rigging and detailed computational fluid dynamics (CFD) simulations. Therefore, in the initial design phase, a couple of preliminary designs are conducted to establish initial values for combustor performance and geometric characteristics. In these preliminary designs, various one-dimensional models using analytical and empirical formulations may be used. One of the disadvantages of existing models is that they are typically geometric dependant, i.e. they apply only to the geometry they are derived for. Therefore the need for a more versatile design tool exists. In this work, which constitutes the first step in the development of such a versatile design tool, a single equation-set network simulation model to describe both steady state compressible and incompressible isothermal flow is developed. The continuity and momentum equations are solved through a hybrid type network model analogy which makes use of the SIMPLE pressure correction methodology. The code has the capability to efficiently compute flow through elements where the loss factor K is highly flow dependant and accurately describes variable area duct flow in the case of incompressible flow. The latter includes ducts with discontinuously varying flow sectional areas. Proper treatment of flow related non-linearities, such as flow friction, is facilitated in a natural manner in the proposed methodology. The proposed network method is implemented into a Windows based simulation package with a user interface. The ability of the proposed method to accurately model both compressible and incompressible flow is demonstrated through the analyses of a number of benchmark problems. It will be shown that the proposed methodology yields similar or improved results as compared to other’s work. The proposed method is applied to a research combustor to solve for isothermal flows and flow splits. The predicted flows were in relatively close agreement with measured data as well as detailed CFD analysis. / Dissertation (MEng (Mechanical Engineering))--University of Pretoria, 2005. / Mechanical and Aeronautical Engineering / unrestricted