Relevant bibliographies by topics / Liquid Propellant Rocket

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Academic literature on the topic 'Liquid Propellant Rocket'

Author: Grafiati
Published: 4 June 2021
Last updated: 6 September 2023

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Journal articles on the topic "Liquid Propellant Rocket"

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Ma, Jiaju. "Analysis of the characteristics of rocket propellant." Theoretical and Natural Science 5, no. 1 (May 25, 2023): 490–95. http://dx.doi.org/10.54254/2753-8818/5/20230296.

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Rocket propellant is an important part of the rocket. Solid or liquid propellant will burn in the engine combustion chamber, and then a large amount of high-pressure gas will be generated. High-pressure gas will be ejected from the engine nozzle at a high speed, generating a reaction force on the rocket, so that the rocket will advance in the opposite direction of the gas injection. This paper mainly analyzes the advantages and disadvantages of current propellants and conceives a theoretically feasible propellant selection method. The main method of research is to calculate the theory of each propellant and make diagrams. Some of the research data were based on existing research reports. Each fuel has its own characteristics, they have one aspect of excellent ability, such as heat conduction, high thrust, high reliability. This paper summarizes the characteristics of current propellants and provides a more convenient query for future researches.
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Junqueira Pimont, Lia, Paula Cristina Gomes Fernandes, Luiz Fernando de Araujo Ferrão, Marcio Yuji Nagamachi, and Kamila Pereira Cardoso. "Study on the Mechanical Properties of Solid Composite Propellant Used as a Gas Generator." Journal of Aerospace Technology and Management, no. 1 (January 21, 2020): 7–10. http://dx.doi.org/10.5028/jatm.etmq.65.

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A gas generating propellants are used as initiators of liquid rocket propellants turbopumps and have as desired characteristic a high-volume production of low-temperature gas. In this context, some formulations of composite propellant containing polyurethane (based on liquid hydroxyl-terminated polybutadiene), guanidine nitrate, ammonium perchlorate, and additives were evaluated and characterized in order to verify their potential as gas generator propellant, as well as to evaluate the influence of additives on mechanical properties. The formulations were prepared, analyzed, and tested for mechanical properties.
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Abdelraouf, A. M., O. K. Mahmoud, and M. A. Al-Sanabawy. "Thrust termination of solid rocket motor." Journal of Physics: Conference Series 2299, no. 1 (July 1, 2022): 012018. http://dx.doi.org/10.1088/1742-6596/2299/1/012018.

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Abstract Rocket motors are engines that create the necessary thrust for the rocket motion. There are different types of rocket motors based on the propellant state, such as solid propellant rocket motors, liquid propellant rocket motors, and hybrid propellant rocket motors. One of the biggest disadvantages of solid propellant rocket motors, in comparison to liquid and hybrid propellant rocket motors, is that they are extremely difficult to extinguish, necessitating the use of specific devices. This paper reviews various ways for thrust termination such as fluid injection, rapid increase in throat area, and sudden opening of an additional port at the forward section of the motor, which increases the depressurization rate (dp/dt) required for extinguishing. The rate of depressurization varies depending on propellant components, combustion pressure, and exhaust pressure, and may be investigated using experimental approaches. The change in the critical area for a motor can be predicted by using MATLAB code to ensure the complete extinguishing by decreasing the pressure under the deflagrationlimit with high depressurization rate.
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Czerwińska, Magdalena, and Piotr Prasuła. "STUDY OF THERMO-MECHANICAL PROPERTIES OF AGED HOMOGENEOUS SOLID ROCKET PROPELLANT ACCORDING TO STANAG REQUIREMENTS." PROBLEMY TECHNIKI UZBROJENIA 145, no. 1 (May 15, 2018): 47–63. http://dx.doi.org/10.5604/01.3001.0012.1325.

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Proper operation of rocket motor depends significantly on thermo-mechanical properties of propellant used. For this reason it is important that characteristics of a particular propellant versus the time and temperature pass a thorough investigation to assess its operation at different conditions. The paper illustrates investigations of ageing process influencing thermo-mechanical properties of homogeneous rocket propellant. A selected type of rocket propellant was subjected to accelerated ageing in conditions specified in AOP-48 document to establish in the next step its thermal and mechanical characteristics (between all the temperature of glass transition and decomposition). The ageing of propelling explosives causes the reduction of stabiliser content deciding about thermo-mechanical properties of propellant and for that the percentage of effective stabiliser and its loss were identified by liquid chromatography HPLC. Thermal properties were investigated by differential scanning calorimetry. Thermal analyses were carried out according to STANAG 4515. Mechanical characteristics were tested by dynamic mechanical analysis (DMA) in line with STANAG 4540.
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ZAGANESCU, Nicolae-Florin, Rodica ZAGANESCU, and Constantin-Marcian GHEORGHE. "Wernher Von Braun’s Pioneering Work in Modelling and Testing Liquid-Propellant Rockets." INCAS BULLETIN 14, no. 2 (June 10, 2022): 153–61. http://dx.doi.org/10.13111/2066-8201.2022.14.2.13.

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This paper presents a view on how Dr. Wernher Von Braun laid the basis for realistic modelling and testing liquid-propellants rockets, by his PhD Thesis – a secret document in 1934, which remained classified until 1960. Understanding that better mathematical modelling is needed if these rockets are to become spaceflight vehicles, he clarified in his thesis essential issues like: maximum achievable rocket speed; Laval nozzle thrust gain; polytropic processes in the combustion chamber and nozzle; influence of equilibrium and dissociation reactions; original measurement systems for rockets test stand; engineering solutions adequate for series production of the combustion chamber – reactive nozzle assembly. The thesis provided a theoretical and experimental basis for a new concept of the rocket, having a lightweight structure; low tanks pressure; high-pressure pumps and injectors; low start speed; rocket stabilization by gyroscopic means or by active jet controls; longer engine burning time; higher jet speed. Numerous tests made even with a fully assembled rocket (the “Aggregate-I”), improved mathematical model accuracy (e.g., the maximum achievable altitude predicted for the “Aggregate-II” rocket was confirmed later in-flight tests).
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Yuan, Wen-Li, Lei Zhang, Guo-Hong Tao, Shuang-Long Wang, You Wang, Qiu-Hong Zhu, Guo-Hao Zhang, et al. "Designing high-performance hypergolic propellants based on materials genome." Science Advances 6, no. 49 (December 2020): eabb1899. http://dx.doi.org/10.1126/sciadv.abb1899.

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A new generation of rocket propellants for deep space exploration, ionic liquid propellants, with long endurance and high stability, is attracting more and more attention. However, a major defect of ionic liquid propellants that restricts their application is the inadequate hypergolic reactivity between the fuel and the oxidant, and this defect results in local burnout and accidental explosions during the launch process. We propose a visualization model to show the features of structure, density, thermal stability, and hypergolic activity for estimating propellant performances and their application abilities. This propellant materials genome and visualization model greatly improves the efficiency and quality of developing high-performance propellants, which benefits the discovery of new advanced functional molecules in the field of energetic materials.
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Lawrence, Lovell. "A LIQUID-PROPELLANT ROCKET MOTOR.*." Journal of the American Society for Naval Engineers 58, no. 4 (March 18, 2009): 642–45. http://dx.doi.org/10.1111/j.1559-3584.1946.tb02717.x.

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Cheng, Yuqiang, and Jianjun Wu. "Particle swarm algorithm-based damage-mitigating control law analysis and synthesis for liquid-propellant rocket engine." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 233, no. 10 (October 31, 2018): 3810–18. http://dx.doi.org/10.1177/0954410018806080.

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The damage-mitigating control is a novel technique to ameliorate the reliability and safety of liquid-propellant rocket engines by achieving an optimized trade-off level between overall dynamic performance of the liquid-propellant rocket engine and structural durability of some selected critical damageable components under the condition of no impact on the achievement of the launch and flight mission. Thus, it is needed to be solved for the damage-mitigating control that the global optimization of the best trade-off between the damage of the critical damageable components and the performance of rocket engine. The major challenge should focus on: (i) to construct model of a certain rocket engine system dynamics, critical components structural dynamics, and damage dynamics; (ii) to optimize open loop feed-forward control law based on liquid-propellant rocket engine system dynamic model, structural and damage dynamics model, by using particle swarm optimization algorithm; (iii) to synthesize an intelligent damage-mitigating control system using the optimized open loop control law. In this paper, synthesis procedure of damage mitigation is introduced; structure and damage dynamic model of damageable components are formulated. The results of the simulation computation show that the synthesized control laws are implemented and achieve the effect of damage mitigating for the liquid-propellant rocket engine. It can provide important theoretical and practical value not only for improving the safety and reliability of the liquid-propellant rocket engine, but also for the complex thermo-flow-mechanical systems such as airplane engines, automobile engines, and fossil-fueled power plant because their service life is very critical too.
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Palacz, Tomasz, and Jacek Cieślik. "Experimental Study on the Mass Flow Rate of the Self-Pressurizing Propellants in the Rocket Injector." Aerospace 8, no. 11 (October 26, 2021): 317. http://dx.doi.org/10.3390/aerospace8110317.

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High vapor pressure propellants such as nitrous oxide are widely used in experimental hybrid and liquid rockets as they can be used in a self-pressurization mode, eliminating the need for external pressurization or pumps and simplifying the design of the rocket system. This approach causes the two-phase flow in the feed system and the injector orifices, which cannot be easily modeled and accounted for in the design. A dedicated test stand has been developed to better understand how the two-phase flow of the self-pressurizing propellant impacts the mass flow characteristics, enabling the simulation of the operating conditions in the rocket engine. The injectors have been studied in the range of ΔP. The flow regimes have been identified, which can be predicted by the SPI and HEM models. It has been shown that the two-phase flow quality upstream of the injector may impact the discharge coefficient in the SPI region and the accuracy of the HEM model. It has been found that the transition to the critical flow region depends on the L/D ratio of the injector orifice. A series of conclusions can be drawn from this work to design the rocket injector with a self-pressurizing propellant to better predict the mass flow rate and ensure stable combustion.
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Khan, Tajwali, and Ihtzaz Qamar. "Factors Affecting Characteristic Length of the Combustion Chamber of Liquid Propellant Rocket Engines." July 2019 38, no. 3 (July 1, 2019): 729–44. http://dx.doi.org/10.22581/muet1982.1903.16.

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Optimum characteristic length of the combustion chamber of liquid rocket engine is very important to get higher energy from the liquid propellants. Characteristic length is defined by the time required for complete burning of fuel. Combustion reactions are very fast and combustion is evaporation dependent. This paper proposes fuel droplet evaporation model for liquid propellant rocket engine and discusses the factors which can affect the required size of characteristic length of the combustion chamber based on proposed model. The analysis is performed for low temperature combustion chamber. A computer code based on proposed model is generated, which solve analytical equations to calculate combustion chamber characteristic length under various input conditions. The analysis shows that characteristic length is affected by combustion chamber temperature, pressure, fuel droplet diameter, chamber diameter, mass flow rate of propellants and relative velocity of the droplet in the combustion chamber.
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Dissertations / Theses on the topic "Liquid Propellant Rocket"

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Boysan, Mustafa Emre. "Analysis Of Regenerative Cooling In Liquid Propellant Rocket Engines." Master's thesis, METU, 2008. http://etd.lib.metu.edu.tr/upload/12610190/index.pdf.

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High combustion temperatures and long operation durations require the use of cooling techniques in liquid propellant rocket engines. For high-pressure and high-thrust rocket engines, regenerative cooling is the most preferred cooling method. In regenerative cooling, a coolant flows through passages formed either by constructing the chamber liner from tubes or by milling channels in a solid liner. Traditionally, approximately square cross sectional channels have been used. However, recent studies have shown that by increasing the coolant channel height-to-width aspect ratio and changing the cross sectional area in non-critical regions for heat flux, the rocket combustion chamber gas side wall temperature can be reduced significantly without an increase in the coolant pressure drop. In this study, the regenerative cooling of a liquid propellant rocket engine has been numerically simulated. The engine has been modeled to operate on a LOX/Kerosene mixture at a chamber pressure of 60 bar with 300 kN thrust and kerosene is considered as the coolant. A numerical investigation was performed to determine the effect of different aspect ratio cooling channels and different number of cooling channels on gas-side wall and coolant temperature and pressure drop in cooling channel.
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Miquel, Valentin. "Propellant Feeding System of a Liquid Rocket With Multiple Engines." Thesis, KTH, Rymdteknik, 2020. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-276460.

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Multiengine first stages are the new trend in recent rockets. Reusability and an oxygen/methane based engine complete this picture. ArianeGroup wants to develop its own rocket following these principles. This thesis presents the study of the feeding system for a seven Prometheus engine rocket. Several ways of connecting propellant tanks to engines were proposed and analyzed. Two configurations were selected and studied with more detail. One consists of a main feeding line which is then split in seven secondary lines. The other one adds one rank of pipes to reduce the number of feeding valves. Their performances were assessed according to classic space industry drivers. Furthermore, the impact of the two solutions on the efficiency of the tank was evaluated. CAD drawings and simulation models were made and could be a base for future work if one of the systems is chosen. The study shows that a falcon 9 like feeding system is performant in terms of mass and pressure losses but another cost-effective configuration is possible and gives good results.
Första stegen med flera motorer är den nya trenden i de senaste raketerna. Återanvändbart och en syre och metan-baserad motor kompletterar denna bild. ArianeGroup vill utveckla sin egen raket enligt dessa principer. Denna avhandling presenterar studien av drivmedelsrör för en sju Prometheus-motorraket. Flera sätt att ansluta drivmedelstankar till motorer föreslogs och analyserades. Två konfigurationer valdes ut och studerades mer detaljerat. En består av en huvudlinje som sedan delas upp i sju sekundära linjer som på SpaceX Falcon 9. Den andra lösningen lägger till en rang av rör för att minska antalet ventiler. Deras prestanda utvärderades först enligt klassiska kriterier för rymdindustrin. Dessutom utvärderades de två lösningarnas påverkan på tankens effektivitet. CAD-ritningar och simuleringsmodeller gjordes och kan vara en bas för framtida arbeten om ett av systemen väljs. Studien visar att ett Falcon 9-liknande konfiguration har bättre prestanda när det gäller massa och tryckförluster men en annan kostnadseffektiv konfiguration är möjlig och ger goda resultat.
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St, Germain Brad David. "Technique for the optimization of the powerhead configuration and performance of liquid rocket engines." Diss., Georgia Institute of Technology, 2003. http://hdl.handle.net/1853/13063.

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Pérez, Roca Sergio. "Model-based robust transient control of reusable liquid-propellant rocket engines." Thesis, université Paris-Saclay, 2020. http://www.theses.fr/2020UPASS017.

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La tendance actuelle vers un accès plus abordable à l'espace se traduit par des lanceurs et moteurs réutilisables. Du point de vue du contrôle, ces moteurs fusée à propergol liquide (MFPL) réutilisables impliquent des spécifications de robustesse plus exigeantes que ceux à usage unique, principalement en raison de leurs capacités de redémarrage multiple et de modulation de poussée. Classiquement, le système de contrôle gère les opérations des MFPL autour d'un ensemble fini de points prédéfinis. Cette approche réduit leur domaine de modulation à un intervalle restreint dans lequel ils sont conçus pour être sûrs. De plus, les phases transitoires, qui ont un impact important sur la vie du moteur, ne sont pas exécutées de manière robuste. L’objectif de ce travail est donc de développer une boucle de régulation adaptée à l’ensemble des phases d'opération (transitoire et régime permanent) et robuste aux variations paramétriques internes. Plusieurs blocs ont été développés pour constituer la boucle de régulation : simulation de moteur, génération de référence et contrôleurs. Des simulateurs représentatifs des moteurs à cycle générateur de gaz ont tout d'abord été construits. La modélisation purement thermodynamique du cycle a ensuite été adaptée au contrôle, afin d'obtenir des modèles non-linéaires sous forme d'état. Dans ces modèles, l'influence des entrées de commande continues (ouvertures des vannes) et des entrées discrètes (activation des allumeurs et démarreur) est considérée dans un cadre hybride simplifié. La sous-phase continue du transitoire de démarrage est contrôlée en boucle fermée pour suivre des trajectoires de référence pré-calculées. Outre le démarrage, les scénarios de modulation présentent également un algorithme pour le suivi des états finaux. Une méthode de contrôle à base de modèles, la commande prédictive, a été appliquée de manière linéarisée avec des considérations de robustesse à tous ces scénarios, dans lesquels des contraintes dures doivent être respectées. Le suivi des points de fonctionnement en pression (poussée) et du rapport de mélange dans l'enveloppe de conception est atteint en simulation tout en respectant les contraintes. La robustesse aux variations des paramètres, qui sont identifiés comme prédominants par des analyses, est également démontrée. Ce travail ouvre la voie à la validation expérimentale par des simulations hardware-in-the-loop ou des tests sur banc d'essai
The current trend towards a more affordable access to space is materialising in reusable launchers and engines. From the control perspective, these reusable liquid-propellant rocket engines (LPRE) imply more demanding robustness requirements than expendable ones, mainly due to their multi-restart and thrust-modulation capabilities. Classically, the control system handles LPRE operation at a finite set of predefined points. That approach reduces their throttability domain to a narrow interval in which they are designed to be safe. Moreover, transient phases, which have a great impact on engine life, are not robustly operated. Hence, the goal of this work is to develop a control loop which is adapted to the whole set of operating phases, transient and steady-state, and which is robust to internal parametric variations. Several blocks have been developed to constitute the control loop: engine simulation, reference generation and controllers. First, simulators representative of the gas-generator-cycle engines were built. The purely thermo-fluid-dynamic modelling of the cycle was subsequently adapted to control, obtaining nonlinear state-space models. In these models, the influence of continuous control inputs (valve openings) and of discrete ones (igniters and starter activations) is considered within a simplified hybrid approach. The continuous sub-phase of the start-up transient is feedback controlled to track pre-computed reference trajectories. Beyond the start-up, throttling scenarios also present an end-state-tracking algorithm. A model-based control method, Model Predictive Control, has been applied in a linearised manner with robustness considerations to all these scenarios, in which a set of hard constraints must be respected. Tracking of pressure (thrust) and mixture-ratio operating points within the design envelope is achieved in simulation while respecting constraints. Robustness to variations in the parameters, which are checked to be predominant according to analyses, is also demonstrated. This framework paves the way to experimental validation via hardware-in-the-loop simulations or in test benches
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Sarotte, Camille. "Improvement of monitoring and reconfiguration processes for liquid propellant rocket engine." Thesis, Université Paris-Saclay (ComUE), 2019. http://www.theses.fr/2019SACLS348/document.

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La surveillance et l'amélioration des modes de fonctionnement des systèmes propulsifs des lanceurs représentent des défis majeurs de l'industrie aérospatiale. En effet, une défaillance ou un dysfonctionnement du système propulsif peut avoir un impact significatif pour les clients institutionnels ou privés et entraîner des catastrophes environnementales ou humaines. Des systèmes de gestion de la santé (HMS) pour les moteurs fusée à ergols liquides (LPREs), ont été mis au point pour tenir compte des défis actuels en abordant les questions de sureté et de fiabilité. Leur objectif initial est de détecter les pannes ou dysfonctionnements, de les localiser et de prendre une décision à l’aide de Redlines et de systèmes experts. Cependant, ces méthodes peuvent induire de fausses alarmes ou des non-détections de pannes pouvant être critiques pour la sécurité et la fiabilité des opérations. Ainsi, les travaux actuels visent à éliminer certaines pannes critiques, mais aussi diminuer les arrêts intempestifs. Les données disponibles étant limitées, des méthodes à base de modèles sont essentiellement utilisées. La première tâche consiste à détecter les défaillances de composants et/ou d'instruments à l'aide de méthodes de détection et de localisation de fautes (FDI). Si la faute est considérée comme mineure, des actions de « non-arrêt » sont définies pour maintenir les performances de l'ensemble du système à un niveau proche de celles souhaitées et préserver les conditions de stabilité. Il est donc nécessaire d’effectuer une reconfiguration robuste (incertitudes, perturbations inconnues) du moteur. Les saturations en entrée doivent également être prises en compte dans la conception de la loi de commande, les signaux de commande étant limités en raison des caractéristiques ou performances des actionneurs physiques. Les trois objectifs de cette thèse sont donc : la modélisation des différents sous-systèmes principaux d’un LPRE, le développement d’algorithmes de FDI sur la base des modèles établis et la définition d’un système de reconfiguration du moteur en temps réel pour compenser certains types de pannes. Le système de FDI et Reconfiguration (FDIR) développé sur la base de ces trois objectifs a ensuite été validé à l’aide de simulations avec CARINS (CNES) et du banc d’essai MASCOTTE (CNES/ONERA)
Monitoring and improving the operating modes of launcher propulsion systems are major challenges in the aerospace industry. A failure or malfunction of the propulsion system can have a significant impact for institutional or private customers and results in environmental or human catastrophes. Health Management Systems (HMS) for liquid propellant rocket engines (LPREs), have been developed to take into account the current challenges by addressing safety and reliability issues. Their objective was initially to detect failures or malfunctions, isolate them and take a decision using Redlines and Expert Systems. However, those methods can induce false alarms or undetected failures that can be critical for the operation safety and reliability. Hence, current works aim at eliminating some catastrophic failures but also to mitigate benign shutdowns to non-shutdown actions. Since databases are not always sufficient to use efficiently data-based analysis methods, model-based methods are essentially used. The first task is to detect component and / or instrument failures with Fault Detection and Isolation (FDI) approaches. If the failure is minor, non-shutdown actions must be defined to maintain the overall system current performances close to the desirable ones and preserve stability conditions. For this reason, it is required to perform a robust (uncertainties, unknown disturbances) reconfiguration of the engine. Input saturation should also be considered in the control law design since unlimited control signals are not available due to physical actuators characteristics or performances. The three objectives of this thesis are therefore: the modeling of the different main subsystems of a LPRE, the development of FDI algorithms from the previously developed models and the definition of a real-time engine reconfiguration system to compensate for certain types of failures. The developed FDI and Reconfiguration (FDIR) scheme based on those three objectives has then been validated with the help of simulations with CARINS (CNES) and the MASCOTTE test bench (CNES/ONERA)
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Vasques, Brunno Barreto. "Numerical and experimental study of swirl atomizers for liquid propellant rocket engines." Instituto Tecnológico de Aeronáutica, 2010. http://www.bd.bibl.ita.br/tde_busca/arquivo.php?codArquivo=3056.

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Presented here is the effort concerning the application of Computational Fluid Dynamics (CFD) as a swirl injector analysis tool. A literature review of the major inviscid swirl models is provided along with the viscous correction procedure. The bi-propellant atomizer design process is described and the features of previous designs are also detailed. The study was undertaken by focusing on the 5-[kN]-thrust rocket engine currently in development. Theoretical predictions of the discharge coefficient, spray angle and liquid film thickness were obtained for both the inner and outer swirlers of the core injection elements. The governing equations are solved based on the laminar volume of fluid (VOF) interface capturing method. Results from cold flow experiments and particle image velocimetry (PIV) are compared to the predictions of the swirl models and the numerical results. The laminar VOF model was able to predict the spray angle with reasonable confidence, however, a deviation of 25 % was observed in the mass flow rate and discharge coefficient. Although the laminar VOF model has proven inadequate, it constitutes a good starting point in the procedure needed to assess swirl injector performance.
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Andersson, Erik. "Preliminary design of a small-scale liquid-propellant rocket engine testing platform." Thesis, Luleå tekniska universitet, Rymdteknik, 2019. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-77079.

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Propulsion system testing before mission operation is a fundamental requirement in any project. For both industrial and commercial entities within the space industry, complete system integration into a static test platform for functional and performance testing is an integral step in the system development process. Such a platform - if designed to be relatively safe, uncomplicated and reliable - can be an important tool within academia as well, giving researchers and students a possibility for practical learning and propulsion technology research. In this thesis, a preliminary design for a liquid-propellant rocket engine testing platform to be used primarily for academical purposes is developed. Included in the presented design is a bi-propellant Chemical Propulsion system, gas pressure fed with Gaseous Nitrogen and using Gaseous Oxygen as oxidiser and a 70 % concentrated ethanol-water mixture as fuel. The propellant assembly contains all necessary components for operating the system and performing combustion tests with it, including various types of valves, tanks and sensors. An estimation of the total preliminary cost of selected components is presented as well. Also part of the developed platform design is a small thrust chamber made of copper, water-cooled and theoretically capable of delivering 1000 N of thrust using the selected propellants. A list of operations to be performed before, during and after a complete combustion test is presented at the end of the document, together with a preliminary design of a Digital Control and Instrumentation System software. Due to time limitations, the software could not be implemented in a development program nor tested with simulated parameters as part of this thesis project.
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Cengiz, Kenan. "Development Of An Iterative Method For Liquid-propellant Combustion Chamber Instability Analysis." Master's thesis, METU, 2011. http://etd.lib.metu.edu.tr/upload/12612753/index.pdf.

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Controlling unsteady combustion induced gas flow fluctuations and the resultant motor vibrations is a very significant step in rocket motor design. It occurs when the unsteady heat release due to combustion happens to feed the acoustic oscillations of the closed duct forming a feed-back system. The resultant vibrations concerned may even lead to total failure of the rocket system unless analysed and tested thoroughly. This thesis aims developing a linear numerical analysis method for the growth rate of instabilities and possible mode shape of a liquid-propelled chamber geometry. In particular, A 3-D Helmholtz code, utilizing Culicks spatial averaging linear iterative method, is developed to find the form of deformed mode shapes iteratively to obtain possible effects of heat source and impedance boundary conditions. The natural mode shape phase is solved through finite volume discretization and the open-source eigenvalue extractor, ARPACK, and its parallel implementation PARPACK. The iterative method is particularly used for analyzing the geometries with complex shapes and essentially for disturbances of small magnitudes to natural mode shapes. The developed tools are tested via two simple cases, a duct with inactive flame and a Rijke tube, used as validation cases for the code particularly with only boundary contribution and heat contribution respectively. A sample 2-D and 3-D liquid-propelled combustion chamber is also analysed with heat sources. After comparing with the expected values, it is eventually proved that the method should be only used for determining the modes instability analysis, as to whether it keeps vibrating or decays. The methodology described can be used as a preliminary design tool for the design of liquid-propellant rocket engine combustors, rapidly revealing only the onset of instabilities.
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Masquelet, Matthieu M. "Simulations of a Sub-scale Liquid Rocket Engine: Transient Heat Transfer in a Real Gas Environment." Thesis, Available online, Georgia Institute of Technology, 2006, 2006. http://etd.gatech.edu/theses/available/etd-11102006-082702/.

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Masquelet, Matthieu Marc. "Large-eddy simulations of high-pressure shear coaxial flows relevant for H2/O2 rocket engines." Diss., Georgia Institute of Technology, 2013. http://hdl.handle.net/1853/47522.

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The understanding and prediction of transient phenomena inside Liquid Rocket Engines (LREs) have been very difficult because of the many challenges posed by the conditions inside the combustion chamber. This is especially true for injectors involving liquid oxygen LOX and gaseous hydrogen GH₂. A wide range of length scales needs to be captured from high-pressure flame thicknesses of a few microns to the length of the chamber of the order of a meter. A wide range of time scales needs to be captured, again from the very small timescales involved in hydrogen chemistry to low-frequency longitudinal acoustics in the chamber. A wide range of densities needs to be captured, from the cryogenic liquid oxygen to the very hot and light combustion products. A wide range of flow speeds needs to be captured, from the incompressible liquid oxygen jet to the supersonic nozzle. Whether one desires to study these issues numerically or experimentally, they combine to make simulations and measurements very difficult whereas reliable and accurate data are required to understand the complex physics at stake. This thesis focuses on the numerical simulations of flows relevant to LRE applications using Large Eddy Simulations (LES). It identifies the required features to tackle such complex flows, implements and develops state-of-the-art solutions and apply them to a variety of increasingly difficult problems. More precisely, a multi-species real gas framework is developed inside a conservative, compressible solver that uses a state-of-the-art hybrid scheme to capture at the same time the large density gradients and the turbulent structures that can be found in a high-pressure liquid rocket engine. Particular care is applied to the implementation of the real gas framework with detailed derivations of thermodynamic properties, a modular implementation of select equations of state in the solver. and a new efficient iterative method. Several verification cases are performed to evaluate this implementation and the conservative properties of the solver. It is then validated against laboratory-scaled flows relevant to rocket engines, from a gas-gas reacting injector to a liquid-gas injector under non-reacting and reacting conditions. All the injectors considered contain a single shear coaxial element and the reacting cases only deal with H₂-O₂ systems. A gaseous oyxgen-gaseous hydrogen (GOX-GH₂) shear coaxial injector, typical of a staged combustion engine, is first investigated. Available experimental data is limited to the wall heat flux but extensive comparisons are conducted between three-dimensional and axisymmetric solutions generated by this solver as well as by other state-of-the-art solvers through a NASA validation campaign. It is found that the unsteady and three-dimensional character of LES is critical in capturing physical flow features, even on a relatively coarse grid and using a 7-step mechanism instead of a 21-step mechanism. The predictions of the wall heat flux, the only available data, are not very good and highlight the importance of grid resolution and near-wall models for LES. To perform more quantitative comparisons, a new experimental setup is investigated under both non-reacting and reacting conditions. The main difference with the previous setup, and in fact with most of the other laboratory rigs from the literature, is the presence of a strong co-flow to mimic the surrounding flow of other injecting elements. For the non-reacting case, agreement with the experimental high-speed visualization is very good, both qualitatively and quantitatively but for the reacting case, only poor agreement is obtained, with the numerical flame significantly shorter than the observed one. In both cases, the role of the co-flow and inlet conditions are investigated and highlighted. A validated LES solver should be able to go beyond some experimental constraints and help define the next direction of investigation. For the non-reacting case, a new scaling law is suggested after a review of the existing literature and a new numerical experiment agrees with the prediction of this scaling law. A slightly modified version of this non-reacting setup is also used to investigate and validate the Linear-Eddy Model (LEM), an advanced sub-grid closure model, in real gas flows for the first time. Finally, the structure of the trans-critical flame observed in the reacting case hints at the need for such more advanced turbulent combustion model for this class of flow.
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Books on the topic "Liquid Propellant Rocket"

1

Sutton, George Paul. History of liquid propellant rocket engines. Reston, Va: American Institute of Aeronautics and Astronautics, 2006.

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de Iaco Veris, Alessandro. Fundamental Concepts of Liquid-Propellant Rocket Engines. Cham: Springer International Publishing, 2021. http://dx.doi.org/10.1007/978-3-030-54704-2.

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Turbopumps for liquid rocket engines. [Warrendale, PA: Society of Automotive Engineers, 1992.

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D, Cruit W., Smith A. W, George C. Marshall Space Flight Center., and AIAA/ASME/SAE/ASEE Joint Propulsion Conference (32nd : 1996 : Lake Buena Vista, Fla.), eds. Cold-flow study of hybrid rocket motor flow dynamics. [Huntsville, AL]: NASA Marshall Space Flight Center, 1996.

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United States. National Aeronautics and Space Administration., ed. Propellant injection systems and processes. [Washington, DC: National Aeronautics and Space Administration, 1995.

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Bian tui li ye ti huo jian fa dong ji ji qi kong zhi ji shu: Variable thrust liquid propellant rocket engine and its control techniques. Beijing: Guo fang gong ye chu ban she, 2001.

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1965-, Liu Kun, and Cheng Mousen 1971-, eds. Ye ti huo jian fa dong ji dong li xue li lun yu ying yong. Beijing: Ke xue chu ban she, 2005.

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Bian tui li ye ti huo jian fa dong ji ji qi kong zhi ji shu: Variable thrust liquid propellant rocket engine and its control techniques. Beijing: Guo fang gong ye chu ban she, 2001.

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1965-, Liu Kun, and Cheng Mousen 1971-, eds. Ye ti huo jian fa dong ji dong li xue li lun yu ying yong. Beijing: Ke xue chu ban she, 2005.

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Ye ti huo jian tui jin zeng ya shu song xi tong: Liquid rocket propellant and pressurization feed systems. Beijing: Guo fang gong ye chu ban she, 2007.

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Book chapters on the topic "Liquid Propellant Rocket"

1

Greatrix, David R. "Liquid-Propellant Rocket Engines." In Powered Flight, 381–415. London: Springer London, 2012. http://dx.doi.org/10.1007/978-1-4471-2485-6_11.

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Mishra, D. P. "Liquid-Propellant Rocket Engines." In Fundamentals of Rocket Propulsion, 261–307. Boca Raton: CRC Press, 2017.: CRC Press, 2017. http://dx.doi.org/10.1201/9781315175997-8.

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Mishra, D. P. "Liquid-Propellant Injection System." In Fundamentals of Rocket Propulsion, 333–95. Boca Raton: CRC Press, 2017.: CRC Press, 2017. http://dx.doi.org/10.1201/9781315175997-10.

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Peschka, Walter. "Liquid Hydrogen as a Rocket Propellant." In Liquid Hydrogen, 105–15. Vienna: Springer Vienna, 1992. http://dx.doi.org/10.1007/978-3-7091-9126-2_5.

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de Iaco Veris, Alessandro. "Fundamental Concepts on Liquid-Propellant Rocket Engines." In Fundamental Concepts of Liquid-Propellant Rocket Engines, 1–61. Cham: Springer International Publishing, 2020. http://dx.doi.org/10.1007/978-3-030-54704-2_1.

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Kwak, Dochan, and Cetin C. Kiris. "Simulation of a Liquid-Propellant Rocket Engine Subsystem." In Computation of Viscous Incompressible Flows, 139–79. Dordrecht: Springer Netherlands, 2010. http://dx.doi.org/10.1007/978-94-007-0193-9_6.

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de Iaco Veris, Alessandro. "Tanks for Propellants." In Fundamental Concepts of Liquid-Propellant Rocket Engines, 563–656. Cham: Springer International Publishing, 2020. http://dx.doi.org/10.1007/978-3-030-54704-2_6.

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de Iaco Veris, Alessandro. "The Thrust Chamber Assembly." In Fundamental Concepts of Liquid-Propellant Rocket Engines, 63–202. Cham: Springer International Publishing, 2020. http://dx.doi.org/10.1007/978-3-030-54704-2_2.

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de Iaco Veris, Alessandro. "Feed Systems Using Gases Under Pressure." In Fundamental Concepts of Liquid-Propellant Rocket Engines, 203–50. Cham: Springer International Publishing, 2020. http://dx.doi.org/10.1007/978-3-030-54704-2_3.

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de Iaco Veris, Alessandro. "Feed Systems Using Turbo-Pumps." In Fundamental Concepts of Liquid-Propellant Rocket Engines, 251–446. Cham: Springer International Publishing, 2020. http://dx.doi.org/10.1007/978-3-030-54704-2_4.

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Conference papers on the topic "Liquid Propellant Rocket"

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KUZNETSOV, N. "Closed-cycle liquid propellant rocket engines." In 29th Joint Propulsion Conference and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1993. http://dx.doi.org/10.2514/6.1993-1956.

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Jensen, Dale. "Advanced Performance, Liquid Propellant, Rocket Engine." In 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2005. http://dx.doi.org/10.2514/6.2005-4566.

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PAL, S., M. MOSER, H. RYAN, M. FOUST, and R. SANTORO. "Flowfield characteristics in a liquid propellant rocket." In 29th Joint Propulsion Conference and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1993. http://dx.doi.org/10.2514/6.1993-1882.

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DELCHER, R., A. DERGEVORKIAN, and S. BARKHOUDARIAN. "Fiberoptics for liquid propellant rocket engine environments." In 25th Joint Propulsion Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1989. http://dx.doi.org/10.2514/6.1989-2416.

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Ehresman, Charles. "The M.W.Kellogg Company's Liquid Propellant Rocket Venture." In 38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2002. http://dx.doi.org/10.2514/6.2002-3584.

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Sierra, Pablo, Juan M. Tizón, Javier Vilas, and José F. Moral. "Efficient Simulation of Liquid Propellant Rocket Engine Cycle." In 53rd AIAA/SAE/ASEE Joint Propulsion Conference. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2017. http://dx.doi.org/10.2514/6.2017-5007.

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Tizón, Juan M., and Alberto Roman. "A Mass Model for Liquid Propellant Rocket Engines." In 53rd AIAA/SAE/ASEE Joint Propulsion Conference. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2017. http://dx.doi.org/10.2514/6.2017-5010.

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TUCKER, E., B. DEHOFF, and R. MCAMIS. "Liquid-propellant rocket engine testing and analysis capabilities." In 29th Joint Propulsion Conference and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1993. http://dx.doi.org/10.2514/6.1993-1861.

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Delcher, Ray C., Doug K. Dinnsen, and S. Barkhoudarian. "Fiber optics in liquid propellant rocket engine environments." In Microlithography '91, San Jose,CA, edited by Norris E. Lewis and Emery L. Moore. SPIE, 1991. http://dx.doi.org/10.1117/12.24819.

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Demiyanenko, Yuri, Anatoly Dmitrenko, Vladimir Rachuk, Alexander Shostak, Alan Minick, Rod Bracken, and Mark Buser. "Single-Shaft Turbopumps in Liquid Propellant Rocket Engines." In 42nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2006. http://dx.doi.org/10.2514/6.2006-4377.

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Reports on the topic "Liquid Propellant Rocket"

1

Williams, Forman A. Fundamental Studies of Liquid-Propellant Rocket Combustion. Fort Belvoir, VA: Defense Technical Information Center, June 2004. http://dx.doi.org/10.21236/ada425218.

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Williams, Forman A. Combustion Processes and Instabilities in Liquid-Propellant Rocket Engines. Fort Belvoir, VA: Defense Technical Information Center, January 2004. http://dx.doi.org/10.21236/ada420091.

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Yang, Vigor. Liquid-Propellant Rocket Engine Injector Dynamics and Combustion Processes at Supercritical Conditions. Fort Belvoir, VA: Defense Technical Information Center, November 2004. http://dx.doi.org/10.21236/ada428947.

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Williams, F. A. Fundamentals of Acoustic Instabilities in Liquid Propellant Rockets. Fort Belvoir, VA: Defense Technical Information Center, April 1992. http://dx.doi.org/10.21236/ada280446.

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Williams, F. A. Fundamental of Acoustic Instabilities in Liquid-Propellant Rockets. Fort Belvoir, VA: Defense Technical Information Center, April 1997. http://dx.doi.org/10.21236/ada329657.

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Williams, Forman A. Fundamentals of Acoustic Instabilities in Liquid Propellant Rockets Under Transcritical Conditions. Fort Belvoir, VA: Defense Technical Information Center, December 1999. http://dx.doi.org/10.21236/ada372407.

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