<p dir="ltr">A fast model of a Rotating Detonation Combustor (RDC) is developed based on the Method of Characteristics (MOC). The model provides a CFD-like solution of an unwrapped 2D RDC flow field in under 10 seconds with similar fidelity as 2D Reacting URANS simulations. Parametric studies are conducted using the simplified model, and the trends are analyzed to gain insight into the underlying physics of rotating detonation combustors. A methodology to assess the performance of operation with multiple waves is presented. The main effect of increasing waves is found to be the increase in the exit Mach number of the combustion chamber. The design process of a mixer component is also presented. The mixer lies downstream of a channel-cooled RDC with subsonic exit and upstream of a Rolls-Royce M250 helicopter engine in open-loop configuration. The mixer dilutes the RDC exhaust with approximately 250% air to condition the flow for the M250 turbine at steady state operation, while also acting as an isolator with a choked throat to prevent back propagation of pressure waves. The mixer aerodynamic design was completed using 2D axisymmetric RANS simulations, and the mechanical design was evaluated using Ansys Mechanical FEA and was found to be able to survive the high thermal stresses present both during the transient heating and steady state operating condition.</p>

<p dir="ltr">Improving launch vehicle and satellite propulsion system performance directly correlates to the delivery of more mass (or quantity) on orbit from launch vehicles, longer duration satellite missions, and longer ranges for missiles/interceptors. Alternative propulsion devices such as rotating detonation engines (RDEs) offer the potential for significant performance gains but their operability has only been demonstrated on “battle hardened” laboratory devices for rocket applications. The objective of this research was to demonstrate cooling and structural approaches that mature rotating detonation rocket engines (RDREs) to flight like maturation levels.</p><p dir="ltr">Multiple 1.6”/4.1 cm diameter RDE combustors were designed, fabricated, and tested. The RDE tested the most accumulated 309 seconds of hot fire testing and 118 starts/shutdowns. Long duration testing was completed to characterize heat flux and high cycle fatigue (HCF) loading. Large quantities of short duration tests were completed to evaluate thermal cycling impacts to the combustor structure and evaluate low cycle fatigue (LCF) loading. The hardware experienced 118 LCF loadings on the combustor cooling passages, equivalent to the amount of thermal cycle starts and shutdowns. An endurance test was completed at 60 seconds in duration, demonstrating operation well beyond thermal steady state. Additionally, ~3.7 million HCF loadings were placed on the combustor cooling passages, equivalent to the approximate amount of detonation wave passes present for all of the WC 2.0 testing.</p><p dir="ltr">Predicted operating pressures ranged from 5 to 15 atm. The highest-pressure conditions resulted in hot gas wall temperatures exceeding 1000°F on the outerbody of the combustor and injector face temperatures peaking at 350°F. Water calorimetry was used to compute heat fluxes, which were then compared to traditional rocket engine throat level heat fluxes calculated using Bartz equations under average operating conditions. The outerbody heat fluxes reached up to 3.7 kW/cm², while injector face heat fluxes reached a maximum of 1.6 kW/cm². When compared to Bartz throat level values, the outer-body heat fluxes varied from 0.9 to 1.6 times the throat level values, and injector heat fluxes ranged from 0.3 to 0.5 times the throat level values.</p><p dir="ltr">A combined thermal and pressure loading fatigue assessment was completed that took into consideration mean stresses and cumulative damage from the spectrum of loading events. Traditional rocket combustor life is typically limited by the thermal cycles that can be placed on the cooling channel hot wall. The fatigue analysis results highlight the reduction in available low cycle fatigue life as RDE's experience larger thermal loads when compared to traditional rocket combustors. Low cycle fatigue life will become especially challenging in higher chamber pressure combustors where thermal environments are more extreme, and the ability keep hot wall temperatures within acceptable levels is more challenging.</p><p dir="ltr">The study also highlights that the passing detonation wave provides a high cycle fatigue (HCF) failure mechanism that is not present in traditional rocket combustors. This failure mechanism is the result of the pressure pulse provided by the passing detonation wave causing a variable load on the hot wall. This variable load is applied at frequencies commonly in the 10's of kHz, resulting in large quantities of loading cycles when operated at rocket like durations (>60 sec). This HCF failure mechanism is most impactful at larger chamber pressures where the detonation pressure ratio causes peak pressures to be elevated, resulting in larger cyclic stresses and strains in the hot wall. The results indicate that high chamber pressure combustors may experience HCF life exceedances within seconds of operation.</p>

No description available.

<p dir="ltr">Burst-mode lasers and burst-mode optical parametric oscillators (OPOs) are applied and developed for planar laser induced fluorescence (PLIF) measurements of key species for high-speed combustion measurements. OH-PLIF in the rotating detonation engine was performed for the first time at wave structure visualization in two different planes and was 10 times faster than any other burst mode OH-PLIF measurements at the time. The same system was used to perform another OH-PLIF experiment at 1 MHz for ~200 pulses to compare key features of the detonation wave structure with computational fluid dynamic simulations and a fundamental detonation tube experiment. The system was also used for seedless velocity measurements in the exhaust by tracking a pocket of OH with a technique called FLASH. A similar OPO was built, aligned, and tuned to perform 1 MHz NO PLIF in a Mach 10 hypersonic tunnel to visualize second mode instabilities and calculate the frequency in the boundary layer transition of a 7-degree cone. A high-efficiency OPO was developed and characterized utilizing the KTP crystal to provide narrow bandwidth pulses for the fluorescence of multiple species. The OPO was pumped at repetition rates up to 1 MHz and was calculated to have a 1.9% UV efficiency from the fundamental 1064 nm output. This is 3 – 5 times increase in efficiency from previous custom and commercial built OPOs. The OPO was applied to the RDC for OH PLIF in the combustor channel and NO PLIF for injector dynamics and response studies. Lastly, a burst-mode laser was used to perform LII on the post detonation blast flow field to measure explosively generated soot. The data was taken at 1 MHz and compared and corrected with a separate set of experiments and computational simulations.</p>

<p dir="ltr">Rotating detonation rocket engines (RDREs) are a quickly-growing area of research in the field of chemical rocket propulsion. The promise of increased thermodynamic efficiency from the pressure-gain combustion cycle has promoted research and development of RDRE technology for an ever-expanding range of systems. Recently, there has been interest in developing RDRE technology for in-space propulsion systems. At the Purdue Altitude Chamber facility, some of the first testing of an RDRE operating with state-of-the-art in-space propellants monomethyl hydrazine (MMH) and mixed oxides of nitrogen (MON) has been performed at simulated altitudes using an experimental 100-lbf-scale thruster from GHKN Engineering, LLC. This work presents the development, modeling, and operation of the test facility at the Maurice J. Zucrow Laboratories for the simulated-altitude testing of an RDRE using hypergolic propellants. The first chapter presents a brief history of the Purdue Altitude Chamber Facility and background information about rotating detonation rocket engines and of supersonic exhaust diffusers, which are commonly-used devices in simulated-altitude testing of rocket engines. The second chapter presents an overview of the test facility and its design, including the vacuum system, the propellant feed system, and the cold flow test stand used to support the test program. The propellant feed system is analyzed in detail to assess the accuracy of theory in predicting the behavior of the feed system in propellant delivery and measurement. The third chapter presents a computational study of the supersonic exhaust diffuser and comparison with test data. Computational methods are developed for the prediction of diffuser performance with RDREs and other annular combustion devices. While rotating detonation has not yet been observed in testing, hot-fire results corroborate the predicted behavior of an annular combustor with a second-throat exhaust diffuser. Operating an RDRE at simulated altitude and with hypergolic propellants MMH and MON represents a novel field of research in detonation-based combustion devices. The development of the test facility for this work at Zucrow Labs is a significant step towards further developing RDRE technology for in-space propulsion applications.</p>

<p dir="ltr">Interest in using detonation-based combustion cycles for use propulsion and power generation has gained considerable attention in the last 10 years or so. The rotating detonation engine (RDE), in particular, has garnered the most attention as a possible replacement for current generation combustion systems. RDEs are continuous flow devices that typically operate in a non-premixed fashion. Reactants are injected into an annular combustion chamber that is usually several millimeters wide. One or more detonation waves propagate azimuthally around the annulus, consuming the reactants. The products then expand out of the combustor where it can produce thrust or be passed into a turbine. The detonation wave front in RDEs travel at speeds between 1-3 km/s which poses additional complexity beyond traditional combustors. There are large gaps in the research community for RDEs that use one or more liquid based propellants. Questions regarding liquid breakup, atomization, breakup, recovery all remain unanswered both experimentally and numerically. This work seeks to understand these fundamental physical phenomena that drive these devices by applying advanced, high-speed laser and other optical diagnostics. </p><p dir="ltr"> A 120 mm nominal diameter rotating detonation combustor that operates on non-premixed hydrogen-air was modified to remove a hydrogen orifice and was replaced with a single liquid fuel injector. This simple, yet important, modification enables the study of a one-way coupling between a liquid fuel jet and a detonation wave at relevant spatio-temporal scales. Planar laser-induced fluorescence was performed at rates up to 1 MHz to quantify the quasi-steady jet dynamics and the recovery behavior of the single liquid jet. Long-duration PLIF imaging lasting 30-40 detonation periods at 300 kHz was also performed for statistical significance. A diesel liquid-in-crossflow injector was observed to breakup or be removed from the PLIF plane within only a few microseconds. After the detonation wave passes through the spray there is a significant dwell period can last between 20-40% of the detonation period before the new fuel is issued into the channel. The quasi-steady liquid jet trajectory was also compared to a jet-in-crossflow from literature and there is decent agreement in the jet near-field. </p><p dir="ltr"> The same hardware scheme with a different liquid fuel injector was tested in conjunction with an alternative imagine scheme. The first technique was able to capture details in the radial-axial plane but could not resolve any motion in the azimuthal direction. A volume-based illumination scheme was used for LIF to image a liquid fuel jet in the azimuthal-axial plane. For this experiment the location of the liquid fuel jet was moved into a different position and as a result experiences significantly different behavior than the jet in crossflow. The breakup and evaporation process takes place over a much longer period of time and there is no pause of liquid fuel injection. Similarly, LIF was performed at 300 kHz for 30 detonation cycles to enable sadistically quantification and phase averaging. Filtered OH* and CH* chemiluminescence imaging was also performed over the same field of view as the LIF imaging. Estimation of the velocity field was calculated using optical flow from the Jet-A LIF images. The velocity results agree well with the recovery analysis from the PLIF measurements.</p><p dir="ltr"> Using the same liquid fuel injection scheme, Jet-A droplet diameter and velocity was measured <i>in-situ</i> during a hot-fire experiment using phase Doppler interferometry (PDI). Although a point technique, PDI was used to measure thousands of droplets during a single test at multiple locations and with multiple conditions. As a means of comparison, cold flow experiments were performed with water in the exit plume. Droplet diameters were measured between 1-20 µs in both cases. PDI results were compared with the optical flow results and there is agreement in median velocities and some differences in the minimum and maximum velocity values. Possible sources of error in the diameter measurement are discussed as well.</p>