PhD, University of Cincinnati, 2011, Engineering and Applied Science: Aerospace Engineering

Combustion at high pressure in applications such as rocket engines and gas turbine engines commonly experience destructive combustion instabilities. These instabilities results from interactions between combustion heat release, fluid mechanics and acoustics. This research explores the significant affect of unstable fluid mechanics processes in augmenting unstable periodic combustion heat release. The frequency of the unstable heat release may shift to match one of the combustors natural acoustic frequencies which then can result in significant energy exchange from chemical to acoustic energy resulting in thermoacoustic instability. The mechanisms of the fluid mechanics in coupling combustion to acoustics are very broad with many varying mechanisms explained in detail in the first chapter. Significant effort is made in understanding these mechanisms in this research in order to find commonalities, useful for mitigating multiple instability mechanisms. The complexity of combustion instabilities makes mitigation of combustion instabilities very difficult as few mitigation methods have historically proven to be very effective for broad ranges of combustion instabilities. This research identifies turbulence intensity near the forward stagnation point and movement of the forward stagnation point as a common link in what would otherwise appear to be very different instabilities. The most common method of stabilization of both premixed and diffusion flame combustion is through the introduction of swirl. Reverse flow along the centerline is introduced to transport heat and chemically active combustion products back upstream to sustain combustion. This research develops methods to suppress the movement of the forward stagnation point without suppressing the development of the vortex breakdown process which is critical to the transport of heat and reactive species necessary for flame stabilization. These methods are useful in suppressing the local turbulence at the forward s (open full item for complete abstract)

Committee: Ephraim Gutmark PhDDSc (Committee Chair); Shaaban Abdallah PhD (Committee Member); Jeffrey Kastner PhD (Committee Member); Prem Khosla PhD (Committee Member) Subjects: Aerospace Materials

PhD, University of Cincinnati, 2024, Engineering and Applied Science: Aerospace Engineering

A rigorous literature survey on combustion instability, acoustic modeling, dynamic flame modeling, and reduced-order modeling (ROM) was presented to acknowledge the difficulty of accurately predicting combustion instability for a lean, fully premixed combustor operating at speed beyond 5% of the speed of sound while producing significantly long flame. The objective is to validate the feasibility of predicting thermoacoustic instability using ROM for changes in boundary condition (exit blockage ratio) and flow rate on a combustor that produces a non-compact flame. A revised ROM was derived with convective flow rather than assuming stationary flow to predict combustion instability in the combustor that operates at a speed beyond 5% of the speed of sound. The structure of the ROM was revised to account for spatially distributed heat release rather than assuming a singular compact flame. A proposed end boundary condition model was used to improve the solution of the ROM. A bluff body combustor was tested to establish a combustor that can operate at speeds greater than 5% of the speed of sound. The bluff body combustor produced stable and unstable non-compact flame under the same flow condition (inlet mach number and equivalence ratio) as the combustor configuration changes (combustor length and blockage ratio). The stable flame combustor configuration is used for flame transfer function measurement. In contrast, the unstable flame combustor configuration is used for ROM prediction validation. The exit pressure reflection coefficients were measured for three different blockage ratios (69%, 56%, 0%) without combustion as the temperature, flow rate, and frequency change to validate the end boundary model proposed in the literature. The model used to characterize the acoustic exit boundary condition for the revised ROM was successfully validated using the multiple microphone method downstream of the bluff body flame holder. An inlet Mach number of 0.06-0.13 at an equ (open full item for complete abstract)

Committee: Jongguen Lee Ph.D. (Committee Chair); Kwanwoo Kim Ph.D. (Committee Member); Paul Orkwis Ph.D. (Committee Member); Prashant Khare Ph.D. (Committee Member) Subjects: Aerospace Materials

MS, University of Cincinnati, 2022, Engineering and Applied Science: Aerospace Engineering

The feasibility of prediction of thermoacoustic instability using a wave-based reduced order network model is evaluated for a micromixer combustor whose flame length is about 43% of the combustor length. A wave-based reduced order network model using the linear acoustic theory is formulated in such a way that the impact of various modeling input parameters such as non-compactness of flame, temperature distribution in the combustor and acoustic boundary conditions on the prediction of onset of combustion dynamics and instability frequency be evaluated. Both global and local flame transfer functions are experimentally determined and incorporated in the network model as the source terms to determine if the compact flame assumption is valid for the prediction. The acoustic boundary conditions as well as the actual temperature distribution are experimentally measured. Modeling results with the global or the local flame transfer function as the source term and with the measured flame temperature show that the prediction of the combustion dynamics does not agree with the experimental results when the analytically modeled boundary conditions are used, warranting the correct determination of acoustic boundary conditions in order to predict the onset of instability correctly. Also, the implementation of distributed flame enhances the accuracy of prediction of thermoacoustic instability for the flame presented in this study.

Committee: Jongguen Lee Ph.D. (Committee Member); Prashant Khare Ph.D. (Committee Member); Michael J. Hughes M.S. (Committee Member) Subjects: Aerospace Materials

MS, University of Cincinnati, 2019, Engineering and Applied Science: Aerospace Engineering

This research provides an experimental investigation into the behavior of naturally occurring, fuel-rich, unstable combustion of liquid fuel in a gas turbine combustor. Testing was done using an acoustically isolated single nozzle combustion test rig using liquid Jet-A fuel, capable of operating under unstable combustion conditions. A test matrix was established to identify operating conditions that incited high combustion instability in fuel-rich conditions. The pressure and flame intensity emissions were analyzed across the test matrix in both time and frequency domains. The Rayleigh Criterion was applied to identify the presence of thermo-acoustic instabilities using combustor pressure and flame emission fluctuations. The flame response was established to further characterize the flame dynamics using using the combustor velocity and flame emission fluctuations. Periodic flame structure and behavior was determined using a high speed camera and compared to the convection time delays calculated via flame emission measurements. The investigation concluded that for cases with sufficiently high pressure fluctuations, the Rayleigh Criterion identified thermo-acoustic instability in fuel-rich combustion. The high-speed video and flame emission measurements revealed the spatial and temporal characteristics of emissions in the OH*-band, indicating that it serves as the best approximation for heat release in a fuel-rich flame. Additionally, the high speed camera video and flame emission measurements indicated that CH*-band emissions convectively lag OH*-band emissions.

Committee: Jongguen Lee Ph.D. (Committee Chair); San-Mou Jeng Ph.D. (Committee Member); Kwanwoo Kim Ph.D. (Committee Member) Subjects: Aerospace Materials

MS, University of Cincinnati, 2019, Engineering and Applied Science: Aerospace Engineering

The current research investigates the spectral characteristics of combustion noise from a non-premixed methane fueled combustor. Sound pressure measurements were recorded using a near field microphone array. The noise emission was studied for different operating and geometric parameters. The operating parameters investigated in this research include the equivalence ratio and the inlet mass flow rate. The geometric parameters examined here include the flame tube confinement ratio and the swirl nozzle type. The noise spectra were always modulated by peaks. COMSOL eigen frequency analysis was used to relate the different peaks found in the experimental spectra to the corresponding acoustic eigen modes of the combustion system. The low emission air-blast type swirl nozzle produced lower noise levels compared to the pressure atomized swirl nozzle for all stable operating conditions. This research also examines the stability characteristics of the low emission nozzle for different equivalence ratios. The system exhibited thermoacoustic oscillations for higher equivalence ratios. The unstable combustion oscillations were found to couple with an eigen mode of the upstream plenum. High speed OH* imaging was done to capture the OH* fluctuations in the reaction zone for both the unstable and stable conditions. POD and DMD analysis on the flame images revealed the presence of an axial mode associated with the instability and a helical mode associated with PVC. The PVC was dominant at an equivalence ratio close to the lean blowout limit. The axisymmetric thermoacoustic mode was dominant at higher equivalence ratios. Vortex shedding was identified as the driving mechanism for the thermoacoustic instability. Adding a plate with an orifice of one nozzle diameter at the exit of the combustor pushed the onset of instability to a lower equivalence ratio.

Committee: Ephraim Gutmark Ph.D. (Committee Chair); Shaaban Abdallah Ph.D. (Committee Member); Mark Turner Sc.D. (Committee Member); Rodrigo Villalva Gomez Ph.D. (Committee Member) Subjects: Aerospace Materials

PhD, University of Cincinnati, 2018, Engineering and Applied Science: Aerospace Engineering

Recent years have witnessed a notable increase in endeavors resorted to investigating unsteady combustion/pressure processes that offer a prospective increase in stagnation pressure due to a more efficient combustion of fuel. One such pressure gain combustion (PGC) concept is a rotating detonation combustor (RDC). RDCs make use of a rotating detonation wave that travels circumferentially about a hollow or annular chamber at kilohertz frequencies, continually combusting the supplied reactants without the need for more than one initial ignition event. Due to its simplicity in design, which can be integrated into existing systems' architecture, and the lack of moving mechanical components, RDCs are at the forefront of PGC research. The current dissertation deals with the basic mechanics of these combustors. Specifically, the diverse modes of detonative operation in annular and hollow combustor configurations are experimentally studied, and the variables dictating these modes are extracted. The question of what exactly constitutes a rotating detonation combustor is answered, by “converting” a conventional atmospheric deflagrative hollow combustor into an RDC. Further, based on this demonstration, the numerous kinships between RDC operation and decades of observations pertaining to high frequency combustion instabilities in rocket engines are presented and discussed. It is argued that most of the poorly understood phenomena of high frequency instabilities can be explained by detonation-based physics. Finally, evidence is presented that suggests rotating detonations to be type of near-limit detonation behavior. The findings of this study are proposed to be useful for the three different communities of RDC research, rocket engine instabilities and fundamental detonation physics.

Committee: Ephraim Gutmark Ph.D. (Committee Chair); Shaaban Abdallah Ph.D. (Committee Member); Mark Turner Sc.D. (Committee Member) Subjects: Aerospace Materials

PhD, University of Cincinnati, 2017, Engineering and Applied Science: Aerospace Engineering

A self-excited instability of liquid-fueled swirling combustion in a laboratory scale combustor is investigated in this study.Liquid fuel is injected through a simplex nozzle and mixed with a swirling air flow generated by an in-house designed, radial-radial, counter-rotating swirler. Air is preheated to 400F (477K). Air flow rate is maintained constant during tests while fuel flow rate is varied to achieve different global equivalence ratios. Pressure perturbations were monitored and recorded by two dynamic pressure transducers (Model PCB 116B) located at 4 and 9.5 inches downstream of a dome plate. Global chemiluminescence filtered by a CH* filter is recorded by a photomultiplier tube (PMT). The data of perturbed pressure and instant chemiluminescence emission are saved simultaneously through a high-response-data-acquisition board. Instant flame image indicating spatial chemiluminescence fluctuations is recorded by a high-speed camera. Acoustic boundaries of the combustion chamber's inlet and outlet were first characterized. For the no bias flow case, the swirler impedance exhibits classical compact-element characteristics. When bias flow is presented, the swirler impedance greatly depends on flow Mach number. The temperature dependence of the impedance is captured in the wavenumber. The chamber exit impedances were measured and curve-fitted into correlations. Then the flame response to air modulation is studied at four equivalence ratios. A loudspeaker is used to perturb plenum air in a 22.5 inches (0.5715m) long combustion chamber. The flame dynamics are expressed as flame transfer functions which are defined as the ratio of relative global heat release rate oscillations to the relative velocity fluctuations at the base of the flame, and calculated within the range of tested equivalence ratio. The behavior of the flame dynamics indicates that the combustion system is similar to a damped high-order system. At last, the instability map of the studied swirler is e (open full item for complete abstract)

Committee: San-Mou Jeng Ph.D. (Committee Chair); Jongguen Lee Ph.D. (Committee Member); Asif Syed Ph.D. (Committee Member) Subjects: Aerospace Materials

Doctor of Philosophy, The Ohio State University, 2014, Chemistry

Heat release studies of plasma assisted combustion have been performed in fuel-air mixtures excited by nanosecond dielectric barrier discharges initially at room temperature and maintained at low pressure (~40 – 50 torr). The following topics have been extensively investigated: (i) the applicability of pure O2 broadband Rotational Coherent Anti-Stokes Raman Scattering spectroscopy at very low O2 pressures of ~8 torr or less to obtain rotational temperature, (ii) validation of a proposed low temperature fuel-oxidation kinetics mechanism fully decoupled from NOx chemistry, (iii) characterization of nanosecond pulse discharges in a dielectric barrier discharge cell and a pin-to-pin discharge geometry, and (iv) effect of fuel addition on heat release in a pin-to-pin discharge geometry at low pressure. For the first topic, the applicability of pure O2 broadband Rotational Coherent Anti-Stokes Raman Scattering (RCARS) Spectroscopy at very low O2 partial pressure of ~ 8 torr or less to obtain rotational temperature has been demonstrated. Very good experimental precisions of ~ ± 1 to 2 K has been demonstrated for diffuse and volumetric plasmas excited by a repetitively pulsed nanosecond discharge. It is shown that the electron-multiplication feature of an EMCCD camera increases the signal to noise ratio significantly. For the second topic, the pure O2 RCARS system was applied to the dielectric barrier discharge cell to obtain time-resolved temperature measurements in nanosecond pulse discharges in 20% O2-Ar, H2-O2-Ar and C2-H2-O2-Ar mixtures, initially at room temperature, operated at a high pulse repetition rate of 40 kHz, in plane-to-plane double dielectric barrier geometry at a pressure of 40 Torr. Nitrogen was deliberately excluded from the system so as to decouple NOx chemistry from the plasma fuel-oxidation processes. It was found that a 0-D model predictions for temperature are in very good agreement in the baseline mixture without fuel and the hydrogen (open full item for complete abstract)

Committee: Walter Lempert PhD (Advisor); Anne McCoy PhD (Committee Member); Terry Gustafson PhD (Committee Member) Subjects: Chemistry; Engineering; Physical Chemistry

MS, University of Cincinnati, 2011, Engineering and Applied Science: Aerospace Engineering

The well known criterion for combustion instability is called the Rayleigh's criterion. It indicates that, for combustion instability to occur, the heat release rate (q') and pressure oscillation (p') must be in phase. This thesis describes measurement techniques and study methods for combustion instabilities that occurred in the prototype single annular sector Rich-Burn Quick-Mix Lean-Burn (RQL) combustor on the original (short) and new (long) experimental rig configuration with a focus on q' and p' measurements. A change in the configuration of the combustor rig was necessary in order to acquire more precise measurements of forward- and backward-moving acoustic pressure waves within the rig by mounting pressure transducers on preselected locations of the upstream duct, downstream duct and combustion area. Pressure transducers provided such local pressure behaviors as amplitude and frequency per location, also in addition to transfer functions that allow for the calculation of the acoustic impedance at any location within the combustor rig. A high-speed camera was capable of filming a chemiluminescene image, i.e., the rate of heat release through a quartz window that is mounted on the side of the combustor. Two imaging analysis techniques, Proper Orthogonal Decomposition and Fourier Transformation, were applied to the chemiluminescene image obtained by a high-speed video device. Two different test cases were investigated. Both a high and low fuel-to-air ratio were used for the investigation of the Rayleigh's criterion, which was confirmed by the corresponding q' and p' data sets. Finally, the resonance frequency that agrees with combustion instability was well predicted by utilizing the one-dimensional wave propagation theory and the known geometry of the combustor rig, temperature of fluid, and boundary conditions.

Committee: San-Mou Jeng PhD (Committee Chair); Shanwu Wang PhD (Committee Member); Kelly Cohen PhD (Committee Member); Asif Syed PhD (Committee Member) Subjects: Aerospace Materials

PhD, University of Cincinnati, 2007, Engineering : Aerospace Engineering

This dissertation is a complementary experimental and theoretical investigation of combustion instability and lean blowout (LBO) in dry-low-emission (DLE) gas turbine engines, aiming to understand the fundamental mechanisms and shed light on active combustion control. Combustion instability involves complicated physicochemical processes, and many of the underlying mechanisms remain unknown, despite extensive research in the past several decades. A practical control system must be able to achieve satisfactory control performances in the presence of large uncertainties, large variations, and even unknown system dynamics. Toward this goal, an observer-based controller, capable of attenuating multiple unstable modes with unknown characteristics, is developed. A mechanism suitable for online prediction of the safety margin to the onset of combustion instability is presented, which does not require knowing the unstable frequencies. The shortage of a reliable, high-frequency, proportional fuel actuator is a major technical challenge for active combustion control. A complementary theoretical and experimental study is performed on a pump-style, high-frequency, magnetostrictive fuel actuator. Improvements to the fuel setup have been made according to the model predictions, which have been experimentally shown to be beneficial to combustion instability control. The second part of this dissertation is about modeling, prediction, and control of lean blowout. The experimentally observed, “intensified”, low frequency, near-LBO combustion oscillations have been used as incipient LBO precursors, and are characterized as low-dimension chaotic behavior in the present study. The normalized chemiluminescence RMS and the normalized cumulative duration of LBO precursor events are recommended for LBO prediction in generic gas turbine engines. Linear stability analysis shows that, with decreasing equivalence ratios, a complex conjugate pair of eigenvalues emerges from three negative real (open full item for complete abstract)

Committee: Dr. Ephraim Gutmark (Advisor) Subjects: