The rapidly expanding space sector is at the forefront of innovation. New technologies are been continuously developed to allow more availability of space for a multitude of commercial or scientific goals. The same is especially true for the field of Space propulsion, where the focus is towards more compact and greener solutions, for launchers, satellites and landers. One of the most suitable candidates for chemical propulsion is the use of liquid oxygen in combination with liquid hydrogen, which, however, comes with many drawbacks connected primarily to the low energetic density of liquid hydrogen and the complexity of storing cryogenics. An innovative solution to this challenge comes with the use of Metal oxidation or metal combustion reaction. This implies the use of the reaction between air and metals or between water and metals to generate heat, power and hydrogen. This allows for much easier power generation since metal powders are simple to stock and have a much higher density than hydrogen. Therefore, the process is compact and completely renewable. The technology has undoubted potential for space applications too. The high energy density, the lack of cryogenics, the high availability and the re-usability make this technology suitable for power generation purposes and, in this case, for propulsive purposes. This thesis aims to explore the various applications of metal combustion, with a particular focus on space propulsion applications. The gathered literature will be then used to produce a conceptual design of a novel propulsion system which maximises the benefits of metal combustion.

A technical project alongside the University courses can deepen the understanding and increase the motivation for the subject of choice. As a student, there is often a hurdle to start such a project because of a lack of inspiration. And even after overcoming this, the costs associated with such a project may put students off. With my project I show how a 3rd semester Mechanical Engineering student can design and manufacture a rocket engine with all testing components on a student budget. Cost structure and resource planning are explained in detail. I launched the project in December 2020 and in September 2021 it was presented at the StuFoExpo21. A general curiosity for the topic and a basic understanding of mechanical engineering was sufficient for starting the project. Importantly, I gained the most valuable knowledge during the implementation of the project, through active failure-iteration and reading specialised literature. The project is focussed on the design and manufacturing of a rocket engine and its testing components. A special feature is the cooling jacket of the combustion chamber. It has been 3D printed in the SLUB Makerspace, a facility at TU Dresden. Further work packages of the project were the programming of sensors and control systems, first open-air combustion tests of the injector head, safety checks and a Risk & Safety analysis. The first testing and other preliminary work were performed in collaboration with fellow students. During the entire design and manufacturing process I was in continuous exchange with the research group “Space Transportation” of the Institute of Aerospace Engineering at TU Dresden. Special thanks go to Dipl.-Ing. Jan Sieder-Katzmann and Dipl.-Ing. Maximilian Buchholz for their help during this process. For 2022 I plan a test campaign of the rocket engine to collect sensor data and to perform engine thrust measurements. / Ein selbständiges, fachbezogenes Projekt neben dem eigentlichen Studium kann die Kenntnisse vertiefen und die Motivation für das Studienfach fördern. Oft jedoch fehlt Studenten die Inspiration für ein solches Projekt. Und selbst wenn diese Hürde genommen ist, schrecken die Kosten davon ab zu beginnen. Am Beispiel meines Projektes im Bereich Luft- und Raumfahrt zeige ich, wie man als Maschinenbaustudent im 3. Semester einen Raketenantrieb und alle Testkomponenten mit einem Studentenbudget entwerfen und herstellen kann. Kostenstruktur und Ressourcenplanung werden im Detail erläutert. Das Projekt startete ich im Dezember 2020. Im September 2021 wurde es auf der StuFoExpo21 vorgestellt. Für den Beginn reichten Neugier und ein Grundlagenverständnis im Maschinenbau aus. Die meisten und wichtigsten Wissensbausteine erlernte ich während des Projektes durch aktive Fehleriteration und aus der Fachliteratur. Das Projekt umfasst das Design und die Fertigung eines Raketenantriebs und der dazugehörigen Testkomponenten. Eine Besonderheit, die Kühlkammer-Ummantelung des Triebwerks, wurde unter Nutzung der Ressourcen an der TU Dresden, SLUB, mit einem 3D-Drucker hergestellt. Weitere wichtige Schritte waren die Programmierung der Sensorik und der Steuerungseinheiten für den Test, ein erster offener Injektor Test mit Treibstoffverbrennung, Sicherheitstests und eine „Risk & Safety Analysis“. Tests und Vorbereitungsarbeiten erfolgten in Zusammenarbeit mit Kommilitonen - der gesamte Entwicklungsprozess fand in ständigem Austausch mit der Forschungsgruppe „Raumtransportsysteme“ des Instituts für Luft- und Raumfahrt in Dresden statt. Besonderer Dank gilt Dipl.-Ing. Jan Sieder-Katzmann und Dipl.-Ing. Maximilian Buchholz für ihre wertvollen Hinweise. Für 2022 ist eine Testkampagne des Raketentriebwerks geplant, um Sensorwerte aufzuzeichnen und den Schub des Triebwerks zu messen.

<p dir="ltr">This work studies the kinetics of nitrous oxide decomposition over alumina-based catalysts, specifically at the high temperatures and high nitrous oxide (N2O) concentrations that would be experienced in catalyst beds for monopropellant rocket thruster applications. High- and low- order models are developed to understand the interaction between reaction kinetics and mass transfer in monolith catalyst tubes. However, nitrous oxide decomposition is not observed on monolith catalyst tubes on account of their lower geometric surface area leading to a majority of the gas not coming into contact with the catalyst surface. Pellet-bed catalysts are studied for the remainder of this work, starting from experiments with a constant-volume batch reactor. The batch experiments demonstrate N2O decomposition over catalyst pellets, and a one-dimensional, time-varying model is developed to quantify the reaction rate based on measured temperature and pressure rise from experimental data. The reaction rates predicted by the model are significantly lower than predicted in the literature for the same catalysts. The inaccuracy is attributed to the fact that the model cannot capture N2O decomposition occurring during the first few seconds of filling the batch tube. Additionally, the simplified temperature distribution applied in the model may not be accurate, and obtaining a higher fidelity temperature distribution experimentally would require more advanced diagnostics.</p><p dir="ltr">The final experiment is a conventional flow-through pellet bed reactor which uses infrared spectroscopy to measure the concentration of nitrous oxide in the decomposed gas mixture. The analysis method incorporates uncertainties from infrared measurements and other sources, and initial activity results of a cobalt oxide-on-alumina catalyst are consistent with the literature. Results from additional testing indicate that manganese oxide catalysts are more active than nickel oxide or cobalt oxide catalysts. At weight loadings of ~10%, results indicate that the Arrhenius pre-exponential constant is roughly an order or magnitude greater for manganese oxide catalysts than cobalt or nickel oxide catalysts. The results also indicate hysteresis in catalytic activity of all oxides. Surface area and x-ray diffraction measurements do not reveal any permanent change in the surface area or crystal structure of these catalysts. The findings lead to the conclusion that the temperature and surrounding environment of the catalyst (either nitrous oxide or nitrogen during system purges) cause short-lived changes to the crystal structure of the active phase, leading to the observed hysteresis.</p>

<p dir="ltr">Instantaneous fuel regression rate within a solid fuel ramjet combustor was characterized using X-ray radiography and ultrasonic transducer measurements. Experiments were performed with cylindrical, center-perforated hydroxyl-terminated polybutadiene (HTPB) fuel grains at three mass fluxes (407-561 kg/m2-s) with consistent inlet total temperatures and chamber pressures. Ultrasonic transducer measurements demonstrated changes of web thickness ranging from 7.50-9.85 mm and regression rate measurements ranging from 1.35-1.74 mm/s. Local maxima of change in web thickness due to flow reattachment and erosive burning were consistently measured with the ultrasonic transducers. Changes in port radius on the order of 8-9 mm and regression rates of approximately 1.25 mm/s were deduced from the X-ray radiography images. Structure of the flow reattachment region was evident in measurements from the X-ray radiography images captured near the combustor entrance while images captured at the mid-length of the combustor exhibited more uniform fuel regression profiles. Ultrasonic measurements of change in web thickness were consistently greater in magnitude relative to X-ray radiography measurements. X-ray radiography imaging allowed for the more accurate measurement of fuel regression with the greatest axial spatial resolution while ultrasonic transducer measurements yielded the greatest radial spatial resolution. The change in web thickness calculated with weight-based techniques yielded smaller magnitude measurements of change in web thickness relative to X-ray radiography.</p><p dir="ltr">Time-dependent measurements of web thickness and regression rate along the port of aluminum-loaded and boron carbide-loaded, hydroxyl-terminated polybutadiene (HTPB) fuel grains were measured in a solid fuel ramjet combustor with X-ray radiography. The combustor was operated at three mass flux conditions, ranging from 397-532 kg/m2-s, with consistent chamber pressures and upstream-of-combustor total temperatures of 1313 kPa and 748 K, respectively. A cross-correlation-based edge detection scheme was used to extract the fuel grain edges within X-ray radiography images collected at 15 Hz. Cross-section photographs of the post-combustion fuel grain surfaces exhibited evidence of flow reattachment and large aft-end regression. Aluminized fuel grains exhibited average weight-based regression rates of 1.29-1.48 mm/s, and boron carbide-loaded fuel grains yielded average regression rates of 1.21-1.38 mm/s. Head-end X-ray measurements of change in port radius indicated flow reattachment, particularly for the bottom (theta = 180) edge of the fuel grain. The absolute maximum of change in port radius, which ranged between 8.56-10.31 mm for aluminized fuel grains and 8.22-9.40 mm for boron carbide-containing fuel grains, did not always coincide with the flow reattachment location. Time-averaged regression rate profiles measured with X-ray radiography were relatively uniform along the port axis but smaller in magnitude compared to the weight-based measurements; 1.17-1.35 mm/s for the aluminum-loaded fuel grains and 1.07-1.24 mm/s for the boron carbide-loaded fuel grains. Pre-ignition fuel regression, on the order of 1.5 mm, was determined to be the cause of the over-prediction of regression rate by weight-based measurements compared to X-ray measurements.</p><p dir="ltr">The weight-based average regression rates measured in tests conducted with the axisymmetric solid fuel ramjet test article in its various configurations were compared to quantify the effects of average port air mass flux, bypass air addition, carbon black addition, and metal particle addition on regression rate. Baseline tests without an aft-mixing section or bypass air addition fuel grains containing carbon black yielded a regression rate coefficient of a = 5.33E-2 and an exponent of n = 0.50 for p4 = 1179-1298 kPa. Including an aft-mixing section without bypass air addition yielded regression rates of 0.94-1.04 mm/s due to the increased residence time. Bypass air addition of 14\% bypass ratio reduced the regression rate to 0.83-0.92 mm/s, and 30% bypass ratio reduced the regression rate to 0.80-0.82 mm/s. For otherwise equal tests, adding carbon black to the fuel grain increased the regression rates from 0.76-0.78 mm/s to 0.83-0.92 mm/s (6-21%). Aluminized fuel grains exhibited an increase in regression rate coefficient over the baseline fuel grains from a = 5.33E-2 to a = 6.30E-2 (18%), but the regression rate exponent remained at n = 0.50. Boron carbide (B4C) addition reduced the regression rate exponent to n = 0.46 but increased the regression rate coefficient to a = 7.55E-2; a 42% increase.</p><p dir="ltr">A simplified solid fuel ramjet combustion model which includes (1) turbulent heat convection, (2) radiation, (3) radiation-coupled surface blowing, (4) unsteady sub-surface heat conduction, (5) solid fuel regression, (6) gas-phase combustion, and (7) fuel port hydrodynamics was developed for regression rate prediction over a range of combustor geometries and operating conditions. Turbulent convection was modeled with empirical correlations relating non-dimensional boundary layer transport numbers. Radiative heat transfer was estimated using modified empirical correlations for radiation in a slab hybrid rocket combustor. Hybrid rocket combustion theory was used to model surface blowing. The condensed-phase heat transfer was modeled by solving the unsteady, variable thermophysical property, regressing surface heat equation with an explicit time-integration, finite volume scheme on a non-uniform grid. A general Arrhenius expression was used to estimate the fuel regression rate. Chemical equilibrium calculations for a stoichiometric HTPB/air diffusion flame were used to model the gas-phase combustion. The port gas dynamics were modeled with compressible flow ordinary differential equations. The results of these individual physical processes were examined in detail for a high mass flux (G_air = 561 kg/m2-s) case. Experiments performed in the axisymmetric solid fuel ramjet combustor were simulated in the model, which yielded a lower regression rate versus mass flux exponent of n = 0.39 compared to the experimentally-obtained n = 0.50. A larger parameter sweep of the model yielded a mass flux exponent of n_1 = 0.30, a pressure exponent of n_2 = 0.04, and an inflow total temperature exponent of n_3 = 0.39. These exponents are less than those observed in other works, but the model successfully captured the relative influence of mass flux, chamber pressure, and inflow total temperature.</p><p dir="ltr">A combustion diagnostic consisting of X-ray radiography and thermocouples embedded within the fuel grain was successfully applied and demonstrated in a solid fuel ramjet slab combustor. One representative test condition with an air mass flowrate of 1 kg/s, an upstream-of-combustor static pressure of 560 kPa, and an upstream-of-combustor total temperature of 639 K was examined. Changes in web thickness of approximately 4 mm and steady-state regression rates of 0.35 mm/s were measured at the thermocouple locations. Condensed-phase temperature measurements yielded fuel grain surface temperatures of 820 K and temperature profiles which were compared to theoretical Michelson profiles. The Michelson profile closely matched the thermocouple-measured temperature profile at one axial location. Sub-surface conductive heat fluxes of 0.35 MW/m2, heat fluxes required to vaporize solid fuel of 0.60 MW/m2$, and surface heat fluxes of 0.95 MW/m2$ were estimated using the condensed-phase temperature profiles.</p>

<p dir="ltr">The Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT) was used for a set of experiments studying the effect of isolated roughness elements on boundary-layer transition on a 7° half-angle cone. In quiet flow, the cone was tested at Reynolds numbers of 7.4 × 10e6 /m, 10.2 × 10e6 /m, and 13.0 × 10e6 /m. Tests were also completed at Re = 11.0 × 10e6 /m in noisy flow to examine the effects of freestream noise. The cone was set at both 0° and 6° angle of attack and an isolated, square trip oriented like a diamond with respect to the flow direction was attached before each set of runs. </p><p dir="ltr">Using infrared thermography and pressure transducers, the location of transition onset was estimated for each test. The results followed all expected trends: transition moved upstream as trip height increased, transition occurred earlier at higher freestream Reynolds numbers, and transition was significantly delayed in quiet flow compared to noisy flow. Mean flow solutions were generated to calculate correlation values commonly used to predict transition. Theexperimentaldatawasthenusedinconjunctionwiththesecorrelationvalues to identify a range of critical values that could be used to predict transition behavior. </p><p dir="ltr">Additionally, a z-type schlieren setup was developed for the BAM6QT. Various components were upgraded and standard procedures for aligning the system were developed. A new pulsed laser and high-speed camera were integrated into the system to enable schlieren imaging at up to 1.75M fps. The final configuration allows the schlieren system to be used for various applications with minimal adjustments, and has been utilized in many research projects in the BAM6QT.</p>

<p dir="ltr">The electrification of pneumatic or hydraulic system on aircraft has been shown effective in reducing the fuel burn. Recently, electrifying propulsive loads has attracted a lot of atten- tion to further improve fuel economy. This work focuses on tools to facilitate more electric aircraft at conceptual design stage, particularly assuming a turbo-generator architecture. Specifically, we develop a simulation tool, mimicking SUAVE [1], which allows mission and fuel burn analysis. Major differences from SUAVE include more detailed models of compo- nents in the electric propulsive branch and degrees of freedom to adjust the velocity profile along the entire mission. Based on the simulator, this work further proposes to leverage a gradient-free optimization technique, which optimizes the optimal velocity profile along the entire mission to minimize fuel burn. Simulation results on two aircraft designs, a con- ventional Boeing 737-800 and NASA-STARC-ABL, verify the effectiveness of the proposed tools.</p>

This investigation explores the detonation dynamics of a rotating detonation engine (RDE). Beginning with the general understanding and characteristics of hydrogen and compressed air as a detonation fuel source, this study further develops the experimental approach to incorporating a coal-based fuel mixture in an RDE. There is insufficient prior research investigating the use of coal as part of a fuel mixture and insignificant progress being made to improve thermal efficiency with deflagration. The U.S. Department of Energy's Office of Fossil Energy awarded the Propulsion and Energy Research Laboratory at the University of Central Florida a grant to lead the investigation on the feasibility of using a coal-based fuel mixture to power rotating detonation engines. Through this study, the developmental and experimental understanding of RDEs has been documented, operability maps have been plotted, and the use of a coal-based fuel mixture in an RDE has been explored. The operability of hydrogen and compressed air has been found, a normalization of all operable space has been developed, and there is evidence indicating coal can be used as part of a fuel mixture to detonate an RDE. The study will continue to investigate coal's use in an RDE. As the most abundant fossil fuel on earth, coal is a popular fuel source in deflagrative combustion for electrical power generation. This study investigates how the combustion of coal can become significantly more efficient.

This research was conducted within the framework of the SpaceDrive project funded by the German Aerospace Center to develop propellantless propulsion for interstellar travel. The experiments attempted to measure mass fluctuations predicted by the Mach-effect theory derived from General Relativity and observed through torsion balance measurements by Woodward (2012). The combination of such mass fluctuations with synchronized actuation promises propellantless thrust with a significantly better thrust-to-power ratio than photon sails. Thus, experiments using different electromechanical devices including the piezoelectric Mach-effect thruster as tested by Woodward et al. (2012) were pursued on sensitive thrust balances. The tests were automated, performed in vacuum and included proper electromagnetic shielding, calibrations, and different dummy tests. To obtain appropriate driving conditions for maximum thrust, characterization of the experimental devices involved spectrometry, vibrometry, finite element analysis, and circuit modeling. Driving modes consisted of sweeps, resonance tracking, fixed frequency, and mixed signals. The driving voltage, frequency, stack pre-tension, mounting, and thruster orientation were also varied. Lastly, different amplifier electronics were tested as well, including Woodward’s original equipment. Experiments on the double-pendulum and torsion balances with a resolution of under 10 nN and an accuracy of 88.1 % revealed the presence of force peaks with a maximum amplitude of 100 nN and a drift of up to 500 nN. The forces mainly consisted of switching transients whose signs depended on the device’s orientation. These force transients were also observed in the zero-thrust configurations. No additional thrust was observed above the balance drift, regardless of the driving conditions or devices tested. In addition, finite element and vibrometry analysis revealed that the vibration from the actuator was transmitted to the balance beam. Moreover, simulations using a simple spring-mass model showed that the slower transient effects observed can be reproduced using small amplitude, high-frequency vibrations. Hence, the forces observed can be explained by vibrational artifacts rather than the predicted Mach-effect thrust. Then, centrifugal balance experiments measured the mass of a device subjected to rotation and energy fluctuations, with a precision of up to 10 µg and a high time resolution. The measurements relied on piezoelectric- and strain gauges. Their calibration methods presented limitations in the frequency range of interest, resulting in discrepancies of up to 500 %. However, the tests conducted with capacitive and inductive test devices yielded experimental artifacts about three orders of magnitude below the mass fluctuations of several milligrams predicted by the Mach-effect theory. Although the piezoelectric devices presented more artifacts due to nonlinearity and electromagnetic interaction, all rotation experiments did not show the expected dependence on the rotation frequency. In summary, the search for low thrust and small mass fluctuations consisted of challenging experiments that led to the development of innovative and sensitive instruments, while requiring a careful consideration of experimental artifacts. The results analysis led to the rejection of mass fluctuations and thrusts claimed by Woodward’s Mach-effect theory and experiments. The quest for breakthrough space propulsion must thus continue a different theoretical or experimental path.:List of Figures List of Tables List of Abbreviations List of Variables and Symbols 1. Introduction 1.1 Research Motivation 1.2 Objectives 1.3 Content Overview 1.4 Team Work 2. Literature Review 2.1 Fundamentals of Space Propulsion 2.2 Mach’s Principle 2.3 Woodward’s Mach-effect Theory 2.3.1 Derivation of the Mass Fluctuation Equation 2.3.2 Design of a Mass Fluctuation Thruster 2.4 Woodward-type Experiments 2.5 Force and Transient Mass Measurements 3. Electromechanical Characterization 3.1 Piezoelectric Actuators 3.1.1 Basic Properties 3.1.2 Actuator Design 3.1.3 Mach-effect Thruster Devices 3.1.4 Magnetostrictive Actuator 3.1.5 Numerical Analysis of MET Behavior 3.1.6 Vibrometry Analysis 3.1.7 Impedance Spectroscopy 3.1.8 Circuit Modeling 3.1.9 Predictions 3.2 Electronics 3.2.1 Description 3.2.2 Characterization 3.3 Torsion Balances 3.3.1 Description 3.3.2 Characterization 3.3.3 Simulation 3.4 Double-pendulum Balance 3.4.1 Description 3.4.2 Characterization 3.5 Laboratory Setup 3.5.1 Vacuum Chambers 3.5.2 Software and Test Setup 4. Thrust Balance Experiments 4.1 Torsion Balance I Test Results 4.1.1 Dummy Tests 4.1.2 CU18A 4.1.3 MET03 4.1.4 MET04 4.1.5 Discussion 4.2 Torsion Balance II Test Results 4.2.1 Dummy Tests 4.2.2 MET05 4.2.3 Beam Vibration 4.2.4 Discussion 4.3 Double-pendulum Balance Test Results 4.3.1 Dummy Tests 4.3.2 MET03 4.3.3 Discussion 5. Centrifugal Balance Experiments 5.1 Centrifugal Balance 5.1.1 Description 5.1.2 Centrifugal Devices 5.1.3 Predictions 5.2 Transducer Calibration 5.2.1 Quasi-Static Calibration I 5.2.2 Quasi-Static Calibration II 5.2.3 Dynamic Calibration 5.3 Centrifugal Balance Test Results 5.3.1 Characterization 5.3.2 CD01 5.3.3 CD02 5.3.4 CD03 5.3.5 CD04 5.3.6 CD05 5.4 Discussion & Error Analysis 6 Conclusions 6.1 Research Summary 6.2 Further Research Appendix A Appendix B Bibliography

Cold flow inside a Solid Fueled Ramjet (SFRJ) is simulated using large eddy simulations (LES). A finite element method using a Discontinuous Galerkin bases has been implemented in the open-sourced multi-physics software SU2. Novel LES formulations of the fuel-gas boundary conditions and the heat release due to mixing are obtained using integration by parts over the discontinuous Galerkin bases. The Smagorisnki and wall-adapted subgrid stress model for the scalar variance have been implemented and investigated in twodimensions. Spectral Proper Orthogonal Decomposition is used to analyze CFD results to determine acoustic modes in the ramjet. Peak acoustic frequencies are compared between between numerical and experimental results. Comparisons are made between simulations performed with a 2D axisymmetric domain and full 3D domain. Cold-flow LES simulations show that there are two dominant acoustic modes (St ≡ f/f0 = {3, 18}) in the ramjet and their frequency appears to be invariant to the cavity configuration. The first peak corresponds to a longitudinal mode associated to the chamber fundamental oscillations (with length scale Lc). The second is characterized with radial fluctuations in the mixing chamber and features the maximum chamber radius of the ramjet as its scaling length. Mixed (radial and axial) modes in the intermediate frequency range reveal the effect of a slanted aft wall on the acoustics. Three-dimensional cold flow simulations predicted weak non-symmetric (azimuthal) modes. Hot-flow simulations show a substantial increase in the mean chamber pressure with the addition of the cavity, indicating that it enhances flame-holding in solid-fuel ramjets, in agreement with the experiments. The analysis of the ramjet acoustic modes shows the emergence of low frequency modes in the cavity cases, in agreement with the experiments. Using SPOD, these modes were associated with low frequency breathing of the recirculation region at the nozzle throat. Perturbations are localized in the throat region because of the Mach number pressure scaling. These modes do not seem to affect the pressure fluctuation and thus combustion in the chamber. Together with the emergence of low frequency vortical modes, the cavity supports a decrease in the high-wave number harmonics of the ramjet chamber acoustic mode. These fluctuations are supported by non-linear amplification of the fundamental mode, which is enhanced by the thermo-acoustic coupling. / Master of Science / Novel propulsion designs, such as solid fuel ramjets, present the opportunity of optimizing cavity shapes using additive manufacturing and three-dimensional printing to improve fuelair mixing and lowering the thermo-acoustic feedback. In this work a computational model for solid fuel ramjets is developed and applied to laboratory firing tests performed by Prof Young's group at the advanced propulsion laboratory at Virginia Tech. In order to capture the fine mixing scales a novel discretization of the reactive Navier-Stokes using discontinuous Galerkin bases is implemented in an open source CFD code popular with aerospace graduate students and researchers. Subgrid modelling is implemented to determine the effect of small scales on the PMMA combustion mechanism developed at Virginia Tech. Numerical methods are used to simulate the turbulent flow of air through an axisymmetric cavity.

A successful transition to clean energy hinges on meeting the world's growing energy demand while reducing greenhouse gas emissions. Achieving this will require significant growth in electricity generation from clean and carbon-free energy sources. Several energy providers have already begun the transition from traditional carbon-based fuels to cleaner alternatives, such as hydrogen and hydrogen enriched natural gas. However, there are still many technical challenges that must be addressed when applying these fuels in gas turbines. The application of hydrogen or hydrogen/natural gas blends to advanced class gas turbines, which have higher operating pressures and temperatures has raised concerns about the potential for leakages or fuel sequencing operations where flammable mixtures of fuel and air could auto-ignite. Public information on the auto-ignition of hydrogen in air at atmospheric pressure is well documented. Such data shows the auto-ignition temperature of hydrogen is roughly 100 °C lower than that of methane. Studies also show that as pressure increases, methane's auto-ignition temperature decreases. However, there was insufficient information in the published literature to characterize the influence of pressure on auto-ignition for hydrogen fuel applications. This study describes the test methodology used to evaluate conditions where auto-ignition occurs for various fuel-air mixtures operating at pressures between 1-30 atmospheres and equivalence ratios between 0.2-1.6. Testing was completed with hydrogen, natural gas and blends at various equivalence ratios using a heated volume with multiple reactant delivery methods. Testing was performed for natural gas to validate the test and data collection methods cited in prior published literature. Results indicate that at atmospheric pressures, an increase in hydrogen concentration results in a reduced auto-ignition temperature. However, at 30 atmospheres, the auto-ignition temperature increased with higher hydrogen concentrations. iv Variations of auto-ignition delay times were also observed during the testing and are compared to modeling predictions, providing insight into auto-ignition characteristics.