<p dir="ltr">Rotating detonation combustors are a developing technology with the potential to successfully integrate pressure gain combustion in to modern propulsion devices. Utilization of propagating detonation waves could increase combustion cycle efficiency and reduce combustor size, resulting in an overall increase in system range or payload-carrying capabilities. However, the sensitivity of rotating detonation combustor operation and performance to geometric features, such as injector configuration or chamber length, still needs to be characterized over a wide range of operating conditions. In addition, the hardware configuration that promotes easy ignition into a coherent detonation operating mode does not always maximize combustor performance, especially at low-loss conditions where feedback between chamber and manifold dynamics can exist. Therefore, a rotating detonation combustor with continuously variable geometry capabilities was designed in order to continuously vary any number of hardware design parameters during combustor testing. Not only does the variable geometry combustor enable rapid characterization of operability sensitivity with minimal hardware swaps, it also enables exploration of hysteresis in performance as the combustor is ignited in one configuration and transitioned to a different geometry while maintaining detonative operation.</p><p dir="ltr">The operability of the variable geometry rotating detonation combustor was first characterized with variable fuel injector location. Higher wave speeds were observed at injector locations closest to the oxidizer throat, with decreased wave speed and eventual transition to deflagrative operation occurring at locations farther downstream due to increasing momentum flux ratio. Variation in fuel injection location induced bifurcations in the number of waves, resulting in corresponding changes in wave speed and gross thrust. Hysteresis was observed in these quantities as the direction of injector translation was reversed. Active translation promoted detonative operation of the experiment at conditions and configurations that hitherto operated only in a deflagrative mode with fixed combustor geometry. </p><p dir="ltr">Sensitivity of rotating detonation combustor operation and performance to oxidizer injector pressure drop was characterized using continuous variation of the injector area during combustor operation. Propulsive performance of the combustor was evaluated using thrust and equivalent available pressure, relating them back to reactant supply pressures for assessment of combustor pressure gain. An effective reactant supply pressure was developed in order to combine contributions of both fuel and oxidizer manifold pressures to the total pressure of the system so that pressure gain could be accurately calculated. Pressure gain increased during a test as oxidizer injector area was increased and the corresponding manifold pressure was decreased. At larger injector areas, pressure gain decreased as the operating mode of the combustor transitioned from detonation to deflagration, concomitant with reduction of gross thrust. Modeling of injector recovery time revealed that the injector operated in both choked and unchoked regimes, which was used to explain detonation wave number transitions in the experiment. A broadened range of detonative operability enabled by active variation of combustor geometry resulted in higher performance with lower injector pressure drop.</p><p dir="ltr">Sensitivity of rotating detonation combustor operation and performance to combustor chamber length was characterized using continuous variation of the chamber length during combustor operation. Specific impulse of the combustor remained relatively constant as chamber length was decreased from its maximum values, proving the practicality of efficient packaging for rotating detonation combustors. A limiting chamber length at which combustion could not longer be supported within the chamber was found to exist for every operating condition, resulting in flame blow-out and performance degradation. Modeling of detonation fill height revealed that relatively low specific impulse measurements could be attributed to unburned reactants exiting the chamber, and a more efficient use of reactants was potentially the cause for improved performance at higher mass flow rates as detonation wave number increased and reactant residence time decreased.</p><p dir="ltr">This experiment and the associated analysis has helped further characterize rotating detonation combustor sensitivity to hardware design parameters. The continuously variable geometry capabilities enabled precise identification of geometric parameters that resulted in operating mode transitions. Analysis and modeling of the flow processes within the injector and chamber were used to help explain why these mode transitions occurred, and can be used for future rotating detonation combustor development.</p>

Earlier research has demonstrated that downstream of combustion in a rotating detonation engine, exhaust flow periodically reverses circumferential direction. For small periods, the circumferential flow reaches velocity magnitudes rivaling the bulk flow of exhaust, manifesting as a swirl. The minimization of this swirl is critical to maximizing thrust and engine performance for rocket propulsion. During this study, numerous nozzle contours were iteratively designed and analyzed for losses analytically. Once a nozzle was chosen, further losses were validated through computational fluid dynamics simulations and then tested experimentally. Three different configurations were run with the RDRE: no nozzle, a nozzle without a spike, and a nozzle with a spike. Images of the exhaust quality were recorded using OH* chemiluminescence in high-speed cameras. One camera was used to confirm the existence of a detonation and the frequency of detonation. The second camera is pointed perpendicular to the exhaust flow to capture the quality of exhaust. Quantitative results of the turbulent velocity fluctuations were obtained through particle image velocimetry of the side-imaging frames. All frames in each case were exported and converted to several time-averaged frames whereupon the time-averaged turbulent velocity fluctuation profiles could be compared between cases for swirl attenuation.

When studying pressure-gain combustion and wave rotor combustors, it is vital that any experimental model accurately reflect the real world conditions/applications being studied; this not only confirms previous computational and analytical work, but also provides new insights into how these concepts and devices work in real life. However, mixture stratification can have a noticeable effect on multiple combustion properties, including flame propagation, pressure, ignition time delay, and more, and this is especially true in constant-volume combustion chambers. Because it is beneficial to model wave rotor systems using constant-volume combustors such as what is employed in the IUPUI Combustion and Propulsion Research Laboratory, these stratification effects much be taken into account and reduced if possible. This study sought to find an effective method to reduce stratification in a rectangular constant-volume combustion chamber by means of manual recirculation pump. Spark-ignited flames were first produced in the chamber itself and studied using schlieren and color videography techniques as well as quantitative pressure histories. After determining the pump's effectiveness in reducing stratification, it was next employed when a hot jet of combustion products from a separate combustion chamber was used as an ignition source instead of the spark plug - a process typically employed in real wave rotor combustors. Lastly, the pump was used to study the leakage from the system for future test cases in order to offer further recommendations on how to effectively use the recirculation system. This process found that key properties significant to wave rotor development, such as time ignition delay, were affected by these stratification effects in past studies that did not account for this detail. As such, the pump has been permanently incorporated into the wave rotor model, as stratification is a vital. Additionally, significant fuel leakage is possible during rotational pre-chamber cases, and this should be address before proceeding with such experiments in the future. To combat this, the pump system has been reduced in volume, and suggestions have been provided on how to better seal the main rectangular chamber in the future.

<p>In rotating detonation engines the turbine inlet conditions may be transonic with unprecedented unsteady fluctuations, very different from those in conventional high-pressure turbines. To ensure an acceptable engine performance, the turbine passages must be unchoked at subsonic and started at supersonic conditions. Additionally, to maximize the aerodynamic performance potential, ad-hoc designs are required, suited for the oscillations in Mach number and flow angle. This manuscript focuses on designing and characterizing diffusive turbine vanes that can operate downstream of a transonic rotating detonation combustor.  </p> <p><br></p> <p>First, the phenomenon of unstarting is presented, concentrating on the effect of pressure loss on the accurate prediction of the starting limit. Afterward, a multi-objective optimization with steady Reynolds Averaged Navier Stokes simulations, including the endwall and 3D vane design, is performed. The results are discussed, highlighting the impact of the throat-to-inlet area ratio on the pressure loss and the geometric features of the top-performing designs. Compared to previous  research on stator passages with contoured endwalls, considerable reductions in pressure loss and stator-induced rotor forcing are obtained, with an extended operating range and preserving high turning.  </p> <p><br></p> <p>Subsequently, the influence of the inlet boundary layer thickness on the vane performance is evaluated, inducing remarkable increases in pressure loss and downstream pressure distortion. Employing an optimization with a thicker inlet boundary layer, specific endwall design recommendations are found, providing a notable improvement in both objective functions. The impact of the geometry variations on flow detachment is assessed as well.</p> <p><br></p> <p>Finally, the impact of the inlet flow angle on the vane design is studied through a multi-point, multi-objective optimization with different inlet angles. The effect of incidence on the flow field and vane performance is evaluated first. Then, by comparing the optimized geometries with those optimized for axial inflows, several design guidelines are identified </p>

Les turbomachines actuelles ont atteint un niveau de maturité technique très élevé. De nouvelles architectures reposant sur des cycles thermodynamiques basés sur une combustion à gain de pression, comme la combustion à volume constant (CVC), ont le potentiel d’augmenter leur efficacité. Dans cette étude,une solution qui repose sur l’intégration dans une turbomachine de chambres de combustion à volume constant sans piston (CVCSP) est considérée. Les objectifs de ces travaux de thèse sont doubles : dans un premier temps de développer et de caractériser extensivement un nouveau dispositif (CV2) dédié à la Combustion à volume constant sans piston sur un cas de référence et, dans un second temps, de proposer à travers plusieurs études, une analyse de l’influence de l’aérodynamique et de la dilution sur les processus d’allumage et, plus généralement de combustion. Le dispositif CV2 permet la combustion aérobie en allumage commandé d’un mélange de propane ou de n-décane, injecté directement dans la chambre. Un point de référence est caractérisé en détail via : des mesures de champs de vitesse par PIV, de chimiluminescence pendant la combustion, une analyse 0D développée dans cette étude. La caractérisation détaillée de ce point de référence montre que le dispositif CV2 reproduit correctement une combustion à volume constant turbulente dans un mélange faiblement hétérogène en température et stratifié en composition, et ce sur un nombre de cycles permettant d’établir une convergence statistique raisonnable. Ces diagnostics et analyses sont employés dans 2 cas d’études pour caractériser successivement : l’influence de l’aérodynamique, via une variation de l’instant d’allumage, l’influence des gaz brûlés résiduels sur la combustion en allumage commandé et la stabilité cyclique, via une variation de la pression d’échappement.Dans un fonctionnement sans balayage, on montre que cette variabilité cyclique est liée au premier ordre à la variation de la dilution en gaz brûlé résiduel du mélange et à la vitesse locale. On montre notamment que, pour un mélange donné, il existe une corrélation statistique entre une vitesse statistique limite et la probabilité d’allumage moyenne. Pour représenter l’effet de pression dans un plénum en amont d’une turbine, on réalise une étude paramétrique sur la pression d’échappement. La dilution résultante, croissant avec la pression d’échappement, diminue la vitesse fondamentale de flamme et ralentit donc la combustion. Les niveaux de températures des gaz brûlés résiduels résultent des échanges de chaleur qui ont lieu sur toute la durée du cycle, de l’allumage du cycle N à celui du cycle N+1 suivant. Des extrapolations sur des cycles à température de paroi plus élevée et à échappement plus court montrent que l’adiabaticité du cycle est améliorée (de 20 %) et que l’effet de dilution en température est alors favorable à une vitesse de flamme turbulente qui est alors plus élevée. Un phénomène d’allumage par gaz brûlé résiduel est observé sur certains cycles de combustion. Ce phénomène est caractérisé dans des conditions favorables, i.e. faible richesse (0.66), allumage tardif et cycle plus court. Lors d’un allumage par gaz brûlés résiduels, un noyau de flamme se développe dans les zones présentant des gaz brûlés résiduels chauds et à basse vitesse autour du jet d’admission et se propage ensuite au reste du mélange identiquement à celui qui serait généré par allumage commandé.Ce travail prend place dans le cadre de la chaire industrielle CAPA sur la combustion alternative pour la propulsion aérobie financée par SAFRAN Tech, MBDA et l’ANR. / Current turbomachines have reached a very high level of technical maturity. Thermodynamic cycles based on pressure-gain combustion, such as constant volume combustion (CVC), feature a clear potential for efficiency improvement. The present study considers the integration in a turbomachine of piston-lessCVC chambers. The thesis work is twofold. First, a new experimental setup (CV2) dedicated to cyclic piston-less CVC is developed and thoroughly characterized on a reference operating point. Second, the influence of the aerodynamics and dilution on the processes of ignition and, in a larger sense, on combustion is discussed through dedicated studies. The CV2 device allows for the spark-ignited air-breathing combustion of a mixture of either propane orn-decane, directly injected into the chamber. A reference condition is characterized in details using: PIV velocity field measurements, chemiluminescence of combustion and a 0D modeling of the device. This detailed characterization evidenced that the CV2 combustion chamber successfully replicates, on a number of cycles allowing a reasonable statistical convergence, a turbulent deflagrative constant-volume combustion in a mixture stratified in composition. Those diagnostics and analyses are applied to 2 cases of study to characterize successively : the influence of the aerodynamics, through a variation of the ignition timing, the influence of the residual burnt gases on spark-ignited combustion and the cyclic stability, through a variation of the exhaust backpressure.Operating the device without scavenging of the combustion chamber, we show that the cyclic variability correlates strongly with both the variation of residual burnt gases dilution and the local velocity. Particularly, we show that for a given mixture, a correlation exists between a statistical velocity limit and the average probability of ignition. The effect of a plenum backpressure upstream of a turbine, downstream of the combustion chamber, is simulated by varying the exhaust system backpressure. The resulting dilution, which increases with the exhaust backpressure, diminishes the fundamental flame velocity of the mixture and slows down the combustion. The residual burnt gases temperature results from the integrated heat exchanges that happen during the total cycle duration starting from the end of combustion of cycle N, to the ignition of cycle N+1. Enhanced cycles, with an increased wall temperature and reduced exhaust duration, are extrapolated by 0D analysis. Those cycles evidence a reduction of the cumulated heat exchanges of up to 20 %. The resulting dilutionis more favorable to higher turbulent flame velocity thus to shorter combustion duration. A phenomenon of ignition induced by the residual burnt gases is observed on certain combustion cycles. This phenomenon is characterized in favorable conditions, i.e. fuel-lean equivalence ratio (0.66), late ignition and shortcycles. During an ignition by residual burnt gases, a flame kernel is ignited in areas where the still hot residuals burnt gases meet fresh gases in low-velocity areas around the intake jet. The ignition kernel then propagates to the rest of the mixture in a similar manner as if it was spark-ignited.This work is part of the CAPA Chair research program on Alternative Combustion modes for Air-breathing Propulsion supported by SAFRAN Tech, MBDAFrance and ANR (French National Research Agency).

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>