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State of the art aircraft afterburners employ spray bars to inject fuel and flameholders to stabilize the combustion process. Such afterburner designs significantly increase the length (and thus weight), pressure losses, and observability of the engine. This thesis presents a feasibility study of a compact prime and trigger (PAT) afterburner concept that eliminates the fuel spray bars and flameholders and, thus, eliminates the above-mentioned problems. In this concept, afterburner fuel is injected just upstream or in between the turbine stages. Downstream of the turbine stages, a low power pilot, or trigger , can be used to control the combustion process. The envisioned trigger for the PAT concept is a jet of product gas from ultra-rich hydrocarbon/air combustion that is injected through the afterburner liner. This partial oxidation (POx) gas, which consists mostly of H2, CO, and diluents, rapidly produces radicals and heat that accelerate the autoignition of the primed mixture and, thus, provide an anchor point for the afterburner combustion process. The objective of this research was to demonstrate the feasibility of the PAT concept by showing that (1) combustion of fuel injected within or upstream of turbine stages can occur only downstream of the turbine stages, and (2) the combustion zone is compact, stable and efficient. This was accomplished using two experimental facilities, a developed theoretical model, and Chemkin simulations. The first facility, termed the Afterburner Facility (AF), simulated the bulk flow temperature, velocity and O2 content through a turbojet combustor, turbine stage and afterburner. The second facility, termed the Propane Autoignition Combustor (PAC), was essentially a scaled-down, simplified version of the AF. The developed model was used to predict and interpret the AF results and to study the feasibility of the PAT concept at pressures outside the AF operating range. Finally, the Chemkin simulations were used to study the effect of several POx gas compositions on the afterburner combustion process.

With strict environmental regulations, gas turbine emissions have been heavily constrained. This requires operating conditions wherein thermo-acoustic flame instabilities are prevalent. During this process the combustor acoustics and combustion heat release fluctuations are coupled and can cause severe structural damage to engine components, reduced operability, and inefficiency that eventually increase emissions. In order to develop an engine without these problems, there needs to be a better understanding of the physics behind the coupling mechanisms of this instability. Among the several coupling mechanisms, the “velocity coupling” process is the main focus of this thesis. The majority of literature has treated axisymmetric disturbance fields which are typical of longitudinal acoustic forcing and axisymmetric excitation of ring vortices. Two important non-axisymmetric disturbances are: (1) transverse acoustics, in the case of circumferential modes of a multi-nozzle annular combustor and (2) helical flow disturbances, seen in the case of swirling flow hydrodynamic instabilities. With significantly less analytical treatment of this non-axisymmetric problem, a general framework is developed for three-dimensional swirl-stabilized flame response to non-axisymmetric disturbances. The dynamics are tracked using a level-set based G-equation applicable to infinitely thin flame sheets. For specific assumptions in a linear framework, general solution characteristics are obtained. The results are presented separately for axisymmetric and non-axisymmetric mean flames. The unsteady heat release process leads to an unsteady volume generation at the flame front due to the expansion of gases. This unsteady volume generation leads to sound generation by the flame as a distributed monopole source. A sound generation model is developed where ambient pressure fluctuations are generated by this distributed fluctuating heat release source on the flame surface. The flame response framework is used to provide this local heat release source input. This study has been specifically performed for the helical flow disturbance cases to illustrate the effects different modes have on the generated sound. Results show that the effects on global heat release and sound generation are significantly different. Finally, the prediction from the analytical models is compared with experimental data. First, a two-dimensional bluff-body stabilized flame experiment is used to obtain measurements of both the flow and flame position in time. This enables a local flame response comparison since the data are spatially resolved along the flame. Next, a three-dimensional swirl-stabilized lifted flame experiment is considered. The measured flow data is used as input to the G-equation model and the global flame response is predicted. This is then compared with the corresponding value obtained using global CH* chemilumenescence measurements.

Combustor blowout is a very serious concern in modern land-based and aircraft engine combustors. The ability to sense blowout precursors can provide significant payoffs in engine reliability and life. The objective of this work is to characterize the blowout phenomenon and develop a sensing methodology which can detect and assess the proximity of a combustor to blowout by monitoring its acoustic signature, thus providing early warning before the actual blowout of the combustor. The first part of the work examines the blowout phenomenon in a piloted jet burner. As blowout was approached, the flame detached from one side of the burner and showed increased flame tip fluctuations, resulting in an increase in low frequency acoustics. Work was then focused on swirling combustion systems. Close to blowout, localized extinction/re-ignition events were observed, which manifested as bursts in the acoustic signal. These events increased in number and duration as the combustor approached blowout, resulting an increase in low frequency acoustics. A variety of spectral, wavelet and thresholding based approaches were developed to detect precursors to blowout. The third part of the study focused on a bluff body burner. It characterized the underlying flame dynamics near blowout in greater detail and related it to the observed acoustic emissions. Vorticity was found to play a significant role in the flame dynamics. The flame passed through two distinct stages prior to blowout. The first was associated with momentary strain levels that exceed the flames extinction strain rate, leading to flame holes. The second was due to large scale alteration of the fluid dynamics in the bluff body wake, leading to violent flapping of the flame front and even larger straining of the flame. This led to low frequency acoustic oscillations, of the order of von Karman vortex shedding. This manifested as an abrupt increase in combustion noise spectra at 40-100 Hz very close to blowout. Finally, work was also done to improve the robustness of lean blowout detection by developing integration techniques that combined data from acoustic and optical sensors.

A full Strouhal number thermo-acoustic model is purposed for the feedback control of self excited noise driven combustors. The inclusion of time delays in the volumetric heat release perturbation models create unique behavioral characteristics which are not properly reproduced within current low Strouhal number thermo acoustic models. New analysis tools using probability density functions are introduced which enable exact expressions for the statistics of a time delayed system. Additionally, preexisting tools from applied mathematics and control theory for spectral analysis of time delay systems are introduced to the combustion community. These new analysis tools can be used to extend sensitivity function analysis used in control theory to explain limits to control effectiveness in self-excited combustors. The control effectiveness of self-excited combustors with actuator constraints are found to be most sensitive to the location of non-minimum phase zeros. Modeling the non-minimum phase zeros correctly require accurate volumetric heat release perturbation models. Designs that removes non-minimum phase zeros are more likely to have poles in the right hand complex plane. As a result, unstable combustors are inherently more responsive to feedback control.

Gas turbine reheat thrust augmenters known as afterburners are used to provide additional thrust during emergencies, take off, combat, and in supersonic flight of high-performance aircrafts. During the course of reheat development, the most persistent trouble has been the onset of high frequency combustion instability, also known as screech, invariably followed by rapid mechanical failure. The coupling of acoustic pressure upstream of the flame stabilizer with in-phase heat-release downstream, results in combustion instability by which the amplitude at various resonant modes — longitudinal (buzz — low frequency), tangential or radial (screech — high frequency) – amplifies leading to deterioration of the afterburner components. Various researchers in early 1950s have performed extensive testing on straight jet afterburners, to identify screech frequencies. Theoretical and experimental work at test rig level has been reported in the case of buzz to validate the heat release combustion models. In this work, focus is given to study the high frequency tangential combustion instability by vibro-acoustic software and the tests are conducted on the scaled bypass flow afterburner for confirmation of predicted screech frequencies. The wave equation for the afterburner is solved taking the appropriate geometry of the afterburner and taking into account the factors affecting the stability. Nozzle of the afterburner is taken into account by using the nozzle admittance condition derived for a choked nozzle. Screech liner admittance boundary condition is imposed and the effect on acoustic attenuation is studied. A new combustion model has been proposed for obtaining the heat release rate response function to acoustic oscillations. Acoustic wave – flame interactions involve unsteady kinetic, fluid mechanic and acoustic processes over a large range of time scales. Three types of flow disturbances exist such as : vortical, entropy, and acoustic. In a homogeneous, uniform flow, these three disturbance modes propagate independently in the linear approximation. Unsteady heat release also generates entropy and vorticity disturbances. Since flow is not accelerated in the region of uniform area duct, vortical and entropy disturbances are treated as in significant, as these disturbances are convected out into atmosphere like an open-ended tube, but these are considered in deriving the nozzle admittance condition. Heat release fluctuations that arise due to fluctuating pressure and temperature are taken into consideration. The aim is to provide results on how flames respond to pressure disturbances of different amplitudes and characterised by different length scales. The development of the theory is based on large activation energy asymptotics. One-dimensional conservation equations are used for obtaining the response function for the heat release rate assuming the laminar flamelet model to be valid. The estimates are compared with the published data and deviations are discussed. The normalized acoustic pressure variation in the afterburner is predicted using the models discussed earlier to provide an indication of the resonant modes of the pressure oscillations and the amplification and attenuation of oscillations caused by the various processes. Similar frequency spectrum is also obtained experimentally using a test rig for a range of inlet mean pressures and temperatures with combustion and core and bypass flows simulated, for confirmation of predicted results. Without the heat source only longitudinal acoustic modes are found to be excited in the afterburner test section. With heat release, three additional tangential modes are excited. By the use of eight probes in the circumferential cross section of afterburner it was possible to identify the tangential modes by their respective phase shift in the experiments. Comparison of normalized acoustic pressure and phase with and without the incorporation of perforate liner is made to study the effectiveness of the screech liner in attenuating the amplitude of screech modes. By the analysis, conclusion is drawn about modes that get effectively attenuated with the presence of perforate liner. Parametric study of screech liner porosity factor of 1.5 % has not shown appreciable attenuation. Whereas with 2.5 % porosity significant attenuation is noticed, but with 4 % porosity, the gain is very minimal. Hence, the perforate screech liner with the porosity of 2.5 % is finalized. From the rig runs, first pure screech tangential mode and second screech coupled tangential modes are captured. The theoretical frequencies for first and second tangential modes with their phases are comparable with experimental results. Though third tangential mode is predicted, it was not excited in the experiments. There was certain level of deviation in the prediction of these frequencies, when compared to the experimentally obtained values. For this test section of length to diameter ratio of 5, no radial modes are encountered both in the analysis and experiments in the frequency range of interest. In summary, an acoustic model has been developed for the afterburner combustor, taking into account the combustion response, the screech liner and the nozzle to study the acoustic instability of the afterburner. The model has been validated experimentally for screech frequencies using a model test rig and the results have given sufficient confidence to apply the model for full scale afterburners as a predictive design tool.