The prediction of transient phenomena inside Liquid Rocket Engines (LREs) has not been feasible until now because of the many challenges posed by the operating conditions inside the combustion chamber. Especially, the departure from ideal gas because of the cryogenic injection in a high-pressure chamber is one of the ma jor hurdle for such simula- tions. In order to begin addressing these issue, a real-gas model has been implemented in a massively parallel flow solver. This solver is capable of performing Large-Eddy Simula- tions (LES) in geometrical configurations ranging from an axisymmetric slice to a 3D slice up to a full 3D combustor. We present here the results from an investigation of unsteady combustion inside a small-scale, multi-injectors LRE. Both thermally perfect gas (TPG) and real gas (RG) approaches are evaluated for this LOX-GH2 system. The Peng-Robinson cubic equation of state (PR EoS) is used to account for real gas effects associated with the injection of cryogenic oxygen. Realistic transport properties are computed but simplified chemistry is used in order to achieve a reasonable turnaround time. Results show the impor- tance of the unsteady dynamics of the flow, especially the interaction between the different injectors. The role of the equation of state is assessed and the real gas model, despite a limited zone of application, seems to have a strong influence on the overall chamber behav- ior. Although several features in the simulated results agree well with past experimental observations, the prediction of heat flux using a simplified flux boundary condition is not completely satisfactory. This work also reviews in details the state of our knowledge on supercritical combustion in a coaxial injector configuration, stressing issues where numeri- cal modeling could provide new insights. However, many developments and improvements are required before an LES modeling of such a flow is both feasible and valid. We finally propose a comprehensive roadmap towards the completion of this goal and the possible use of CFD as a design tool for a modern liquid rocket engine.

This work presents a preliminary design of a liquid rocket engine test platform to support research into liquid propulsion systems and rocket engine components, including injectors, ignition systems, combustion chambers and engine cooling systems. The liquid propellants, specifically liquid oxygen and ethanol, are pressure-fed using gaseous nitrogen. The test platform supports engine thrust values up to 1 kN, as well as varying oxidizer/fuel ratios up to 4.0 and varying ethanol concentrations between 70 and 100%. The test platform will integrate with a mobile control centre, which was designed concurrently, and provides remote control of the test procedures and data acquisition of all relevant pressure, temperature, mass flow and thrust data. The propellant feed assembly can support both cold and hot fire testing campaigns and is equipped with numerous safety features including inert gas purge lines, emergency drain lines and emergency shut-down and de-pressurization procedures.

The design of a liquid rocket propulsion system, unlike that of a standalone system, is intertwined with the overall development of a number of associated systems and is influenced by a multitude of conditions and considerations: from the requirements needed to accomplish the mission to the rationalizations involved behind the development of each rocket system and/or component. In my thesis, the preliminary design of a “new generation” 30 kN rocket engine driven by an electric pump feed system and running on liquid methane and liquid oxygen is performed. The propulsion system would be employed on a hypothetical small-lift orbital-class twin-stage rocket to deliver a light payload of about 200 kg into a circular 500 km LEO. Such topics as the selection of bipropellant combinations, the feasibility of electric pump feed systems, design methodologies for thrust chambers, for nozzles in particular, management of the high thermal energy and the selection of compatible wall materials, as well as the design of an injector have been looked comprehensively into. It is realized that methalox is indeed better than both hydrolox (with regard to density impulse) and kerolox (in terms of specific impulse). Besides, a suite of attractive characteristics makes the bipropellant a combination of choice to power rockets of the future. Yet more notably, an electric-pump-fed engine cycle is, under the right circumstances of engine operation, established to outperform both the pressure feed system and the turbopump feed system. With constant advancement in battery technologies, improvement of both power density and energy density to achieve much higher performance is but a matter of time. The adoption of a propulsion system such as ours for a mission objective as outlined above, therefore, is not just viable but unquestionably realistic. Two thrust chamber versions—a sea-level variant for the booster stage and a vacuum-optimized variant for the upper stage—are developed for our rocket. And both the nozzles employ a TOP “thrust optimised parabolic” contour; also, the booster stage comprises a cluster of 9 engines in a parallel burn arrangement. Concerning thermal management, the entirety of the booster-stage thrust chamber implements regenerative cooling (using Inconel 625), whereas the aft of the upper-stage nozzle section implements radiative cooling (with Niobium C-103). Further, the injector faceplate (also of Inconel 625) comprises two concentric patterns of unlike impingement doublet sets: with 80 pairs on the outer ring and 40 pairs on the inner ring. With rational assumptions, our hypothetical launch vehicle is deemed to have a mass of roughly 17200 kg (200 kg of which is the payload) and a delta-v of approximately 9600 m/s—quite within the desirable range of specifications for small-lift orbital-class twin-stage rockets of today.

No description available.

Ces cinquantes dernières années, la majorité des paramètres de conception des moteurs cryotechniques ont été ajustés en l'absence d'une compréhension détaillée des phénomènes de combustion, en raison des limites des diagnostiques expérimentaux et des capacités de calcul. L'objectif de cette thèse est de réaliser des simulations numériques instationnaires d'écoulements réactifs transcritiques de haute fidélité, pour permettre une meilleure compréhension de la dynamique de flamme dans les moteurs cryotechniques et finalement guider leur amélioration. Dans un premier temps, la thermodynamique gaz-réel et son impact sur les schémas numériques sont présentés. Comme la Simulation aux Grandes Echelles (SGE) comporte des équations filtrées, les effets de filtrages induits par la thermodynamique gaz-réel sont ensuite mis en évidence dans une configuration transcritique type et un opérateur de diffusion artificiel, spécifique au gaz réel, est proposé pour lisser les gradients transcritiques en SGE. Dans un deuxième temps, une étude fondamentale du mélange turbulent et de la combustion dans la zone proche-injecteur des moteurs cryotechniques est menée grâce à la Simulation Numérique Directe (SND). Dans le cas non-réactif, les lâchers tourbillonnaires dans le sillage de la lèvre de l’injecteur jouent un rôle majeur dans le mélange turbulent et provoquent la formation de structures en peigne déjà observées expérimentalement dans des conditions similaires. Dans le cas réactif, la flamme reste attachée à la lèvre de l'injecteur, sans extinction locale, et les structures en peigne disparaissent. La structure de flamme est analysée et différents modes de combustion sont identifiés. Enfin, une étude de flamme-jet transcritique H2/O2, accrochée à un injecteur coaxial avec et sans retrait interne, est menée. Les résultats numériques sont d'abord validés par des données expérimentales pour l'injecteur sans retrait. Ensuite, la configuration avec retrait est comparée à la solution de référence sans retrait et à des données experimentales pour observer les effets de ce paramètre de conception sur l'efficacité de combustion. / In the past fifty years, most design parameters of the combustion chamber of Liquid Rocket Engines (LREs) have been adjusted without a detailed understanding of combustion phenomena, because of both limited experimental diagnostics and numerical capabilities. The objective of the present thesis work is to conduct high-fidelity unsteady numerical simulations of transcritical reacting flows, in order to improve the understanding of flame dynamics in LRE, and eventually provide guidelines for their improvement. First real-gas thermodynamics and its impact on numerical schemes are presented. As Large-Eddy Simulation (LES) involves filtered equations, the filtering effects induced by real-gas thermodynamics are then highlighted in a typical 1D transcritical configuration and a specific real-gas artificial dissipation is proposed to smooth transcritical density gradients in LES. Then, a Direct Numerical Simulation (DNS) study of turbulent mixing and combustion in the near-injector region of LREs is conducted. In the non-reacting case, vortex shedding in the wake of the lip of the injector is shown to play a major role in turbulent mixing, and induces the formation of finger-like structures as observed experimentally in similar operating conditions. In the reacting case, the flame is attached to the injector rim without local extinction and the finger-like structures disappear. The flame structure is analyzed and various combustion modes are identified. Finally, a LES study of a transcritical H2/O2 jet flame, issuing from a coaxial injector with and without inner recess, is conducted. Numerical results are first validated against experimental data for the injector without recess. Then, the recessed configuration is compared to the reference solution and to experimental results, to scrutinize the effects of this design parameter on combustion efficiency.

This thesis aims at studying the breakup of swirling liquid sheets discharging from the outer orifice of gas centered swirl coaxial atomizers. Such atomizers are considered as propellant injection systems for semi-cryogenic liquid rocket engines. A gas centered swirl coaxial type atomizer discharges an annular swirling liquid sheet which is atomized by a gaseous jet issuing from the central orifice of the atomizer. The primary objectives of this work were to understand the fluid dynamic interaction process between the outer liquid sheet and the central gas jet and its role on the breakup process of the liquid sheet. Cold flow experiments were carried out by constructing custom made gas centered swirl coaxial atomizers. Two different atomizer configurations with varying swirl effect were studied. The jets were injected into ambient atmospheric air medium with tap water and air as experimental fluids. The flow conditions were described in terms of Weber number (Wel) and Reynolds number (Reg) for liquid sheet and the air jet respectively. Spray images were captured by employing an image acquisition system comprising a high resolution digital camera and a strobe lamp. The captured spray images at different combinations of Wel and Reg were analyzed to extract quantitative measurements of breakup length (Lb), spray cone angle (θs), spray width (SW) and two-dimensional surface profile of liquid sheets. Quantitative analysis of the variation of Lb with Reg with different values of Wel suggested that low inertia liquid sheets undergo an efficient breakup process. High inertia liquid sheets ignore the presence of central air jet at lower values of Reg however undergo air jet breakup at higher values of Reg. Qualitative analysis of experimental observations revealed that the entrainment process, established between the inner surface of the liquid sheet and the boundary of central jet, triggers the air assisted sheet breakup by drawing the liquid sheet closer to the spray axis. The entrainment process may be developing corrugations on the surface of liquid sheet which promotes the production of thick liquid ligaments from the sheet surface. The level of surface corrugations on the liquid sheet, quantified by means of tortuosity of liquid sheet profile, increases with increasing Reg. Limited studies on the effect of variation swirl intensity on the air assisted breakup process of liquid sheets did not show any significant influence for the atomizers examined in the present work.

Combustion instability due to feedback coupling between unsteady heat release and natural acoustic modes can cause catastrophic failure in liquid rocket engines and to predict and prevent these instabilities the mechanisms that drive them must be further elucidated. With this goal in mind, the objective of this thesis was to develop techniques that improve the understanding of the specific underlying physical processes involved in these driving mechanisms. In particular, this work sought to develop a small-scale, optically accessible liquid rocket engine simulator and to apply modern, high-speed diagnostic techniques to characterize the reacting flow and acoustic field within the simulator. Specifically, high-speed (10 kHz), simultaneous data were acquired while the simulator was experiencing a 170 Hz combustion instability using particle image velocimetry, OH planar laser induced fluorescence, CH* chemiluminescence, and dynamic pressure measurements. In addition, this work sought to develop approaches to reduce the large quantities of data acquired, extracting key physical phenomena involved in the driving mechanisms. The initial data reduction approach was chosen based on the fact that the combustion instability problem is often simplified to the point that it can be characterized by an approximately linear constant coefficient system of equations. Consistent with this simplification, the experimental data were analyzed by the dynamic mode decomposition method. The developed approach to apply the dynamic mode decomposition to simultaneously acquired data located a coupled hydrodynamic/combustion/acoustic mode at 1017 Hz. On the other hand, the dynamic mode decomposition's assumed constant operator approach failed to locate any modes of interest near 170 Hz. This led to the development of two new data analysis techniques based on the dynamic mode decomposition and Floquet theory that assume that the experiment is governed by a linear, periodic system of equations. The new periodic-operator data analysis techniques, the Floquet decomposition and the ensemble Floquet decomposition, approximate, from experimental data, the largest moduli Floquet multipliers, which determine the stability of the periodic solution trajectory of the system. The unstable experiment dataset was analyzed with these techniques and the ensemble Floquet decomposition analysis found a large modulus Floquet multiplier and associated mode with a frequency of 169.6 Hz. Furthermore, the approximate Rayleigh criterion indicated that this mode was unstable with respect to combustion instability. Overall, based on the positive finding that the ensemble Floquet decomposition was able to locate an unstable combustion mode at 170 Hz when the operator's time period was set to 1 ms, suggests that the dynamic mode decomposition based 1017 Hz mode parametrically forces the 170 Hz mode, resulting in what could be characterized as a parametric combustion instability.

京都大学 / 新制・課程博士 / 博士(工学) / 甲第24845号 / 工博第5162号 / 新制||工||1986(附属図書館) / 京都大学大学院工学研究科機械理工学専攻 / (主査)教授 平山, 朋子, 教授 松原, 厚, 教授 小森, 雅晴 / 学位規則第4条第1項該当 / Doctor of Philosophy (Engineering) / Kyoto University / DFAM

The noise of a rocket engine is a complex and complicated phenomenon which has been studied for more than half a century [1]. There are many sources to this noise, but due to its great potential impact to the surrounding structures, including the vehicle itself, it is important to have a model to estimate the acoustic environment the engine produces. This estimation can be used for the design of a launch pad or a flame deflector. Such a model was developed and then tested, by measuring the noise levels with six microphones at Rocket Factory Augsburg´s Vertical Test Stand at Esrange Space Centre. Three out of six microphones yielded valuable data. A comparison between the estimated and measured noise was then conducted which showed similar trends. The peak frequency in the estimation was in the order of 1 kHz. A sensitivity study was made to investigate the difference in Sound Power Level (SPL) when the engine and test stand parameters were adjusted. The parameters with the greatest effect on the SPL are the Mach number, thrust, potential core length, and impingement distance. The difference in SPL between the estimation and measured noise is 0-20 dB with a lower difference at lower frequencies and a higher difference at higher frequencies. The difference was higher when comparing the estimation to the test with an overpowered engine, with differences of up to 20 dB higher than the estimation in the upper frequencies. Differences with nominal engine data was up to 15 dB higher than the estimation, constrained to lower frequencies. Above 30% of the peak frequency, the noise was consistently lower than the estimation. The estimation can be concluded to likely be conservative at higher frequencies, further testing or a new estimation is necessary with accurate engine data. / Ljudet från en raketmotor är ett komplext och komplicerat fenomen som har studerats i mer än ett halvt sekel [1].Det finns många källor till det ljudet, men på grund av risken att det skadar omgivande strukturer, inklusive raketen, är det viktigt att ha en modell för att estimera ljudmiljön motorn produceras. Estimeringen kan användas för att designa en uppskjutningsramp eller en flammdeflektor. En sådan modell var utvecklad och testad genom att göra ljudmätningar med sex mikrofoner på Rocket Factory Augsburg´s testanläggning på Esrange Space Centre. Tre av sex mikrofoner gav värdefull data. En jämförelse gjordes mellan det estimerade och uppmätta ljudet vilket visade liknande trender. Toppfrekvensen i estimeringen var i storleksordningen 1 kHz. En känslighetsstudie gjordes för att undersöka skillnaden i ljudnivån (SPL) när motorns och testanläggningens parametrar justerades. Parametrarna med störst påverkan på ljudnivån var Mach numret, drivkraft, längden av flamman och avståndet till deflektorn. Ljudskillnaden mellan det estimerade och uppmätta ljudet var mellan 0-20 dB med mindre skillnad på lägre frekvenser och större skillnad på högre frekvenser. Skillnaden var större vid jämförelse mellan estimering och testet med en kraftigare motor, med skillnader på upp till 20 dB över estimeringen på de högre frekvenserna. Skillnaderna för nominell motordata var upp till 15 dB högre än estimeringen, begränsat till lägre frekvenser. Över 30% av toppferkvensen var ljudet konsekvent lägre än estimeringen. Estimeringen kan sannolikt konstateras vara konservativ på högre frekvenser, ytterligare tester eller estimeringar behövs med exakt motordata.