Thematische Bibliographien / Reusable Launchers

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Auswahl der wissenschaftlichen Literatur zum Thema „Reusable Launchers“

Autor: Grafiati

Veröffentlicht am 23. November 2024

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Zeitschriftenartikel zum Thema "Reusable Launchers"

Baiocco, P., und Ch Bonnal. „Technology demonstration for reusable launchers“. Acta Astronautica 120 (März 2016): 43–58. http://dx.doi.org/10.1016/j.actaastro.2015.11.032.

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Lobanovsky, Yu I. „Efficiency analysis of reusable aerospace launchers“. Aerospace Science and Technology 1, Nr. 1 (Januar 1997): 37–46. http://dx.doi.org/10.1016/s1270-9638(97)90022-5.

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Chelaru, Teodor-Viorel, Valentin Pană und Costin Ene. „Performance Evaluation for Launcher Testing Vehicle“. Aerospace 9, Nr. 9 (09.09.2022): 504. http://dx.doi.org/10.3390/aerospace9090504.

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The paper’s purpose is to present a calculus model for a testing vehicle that can be used to validate guidance, navigation and control systems for reusable launchers in all flight phases. The technical solution is based on a throttleable engine with thrust vectoring control and a reaction control system (RCS) used for roll. For calculus, we will develop a nonlinear model with six degrees of freedom, based on quaternion, extended with nonlinear equations that use pulse modulation in order to control roll. In order to synthesize the controller, we also develop a linear model similar to the launcher model. The paper analyzes two basic scenarios, first with the ascending and the descending flight phases and the second having a horizontal flight interleaved between ascending and descending flight phases, both scenarios being specific for reusable launchers. Based on these scenarios, the paper evaluates some performances of the proposed vehicle, namely flight envelope and guidance accuracy.

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Simplício, Pedro, Andrés Marcos und Samir Bennani. „Guidance of Reusable Launchers: Improving Descent and Landing Performance“. Journal of Guidance, Control, and Dynamics 42, Nr. 10 (Oktober 2019): 2206–19. http://dx.doi.org/10.2514/1.g004155.

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D’Angelo, Salvatore, Edmondo Minisci, Daniele Di Bona und Luciano Guerra. „Optimization Methodology for Ascent Trajectories of Lifting-Body Reusable Launchers“. Journal of Spacecraft and Rockets 37, Nr. 6 (November 2000): 761–67. http://dx.doi.org/10.2514/2.3648.

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Musso, Girolamo, Iara Figueiras, Héléna Goubel, Afonso Gonçalves, Ana Laura Costa, Bruna Ferreira, Lara Azeitona et al. „A Multidisciplinary Optimization Framework for Ecodesign of Reusable Microsatellite Launchers“. Aerospace 11, Nr. 2 (31.01.2024): 126. http://dx.doi.org/10.3390/aerospace11020126.

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The commercial space launch sector is currently undergoing a significant shift, with increasing competition and demand for launch services, as well as growing concerns about the environmental impact of rocket launches. To address these challenges, within the New Space Portugal project scope, a multidisciplinary framework for designing and optimizing new launch vehicles is proposed. Creating a more resilient and responsible space industry can be achieved by combining technological innovation and environmental sustainability, as emphasized by the framework. The main scope of the framework was to couple all the disciplines relevant to the space vehicle design in a modular way. Significant emphasis was placed on the infusion of ecodesign principles, including Life Cycle Assessment (LCA) considerations. Optimization techniques were employed to enhance the design and help designers conduct trade-off studies. In general, this multidisciplinary framework aims to provide a comprehensive approach to designing next-generation launch vehicles that meet the demands of a rapidly changing market while also minimizing their environmental impact. A methodology that leverages the strengths of both genetic and gradient-based algorithms is employed for optimizations with the objectives of maximizing the apogee altitude and minimizing the Global Warming Potential (GWP). Despite only being tested at the moment for sounding rockets, the framework has demonstrated promising results. It has illuminated the potential of this approach, leading to the identification of three optimal designs: one for maximizing the apogee, another for minimizing GWP, and a compromise design that strikes a balance between the two objectives. The outcomes yielded a maximum apogee of 6.41 km, a minimum GWP of 9.06 kg CO2eq, and a balanced compromise design featuring an apogee of 5.75 km and a GWP of 25.64 kg CO2eq.

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Duparcq, J. L., E. Hermant und D. Scherrer. „Turbojet-type engines for the airbreathing propulsion of reusable winged launchers“. Acta Astronautica 29, Nr. 1 (Januar 1993): 41–50. http://dx.doi.org/10.1016/0094-5765(93)90068-8.

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Gulczyński, Mateusz T., Robson H. S. Hahn, Jan C. Deeken und Michael Oschwald. „Turbopump Parametric Modelling and Reliability Assessment for Reusable Rocket Engine Applications“. Aerospace 11, Nr. 10 (02.10.2024): 808. http://dx.doi.org/10.3390/aerospace11100808.

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The development of modern reusable launchers, such as the Themis project with its LOX/LCH4 Prometheus engine, CALLISTO—a reusable VTVL-launcher first-stage demonstrator with a LOX/LH2 RSR2 engine, and SpaceX’s Falcon 9 with its Merlin 1D engine, underscores the need for advanced control algorithms to ensure reliable engine operation. The multi-restart capability of these engines imposes additional requirements for throttling, necessitating an extended controller-validity domain to safely achieve low thrust levels across various operating regimes. This capability also increases the risk of component failure, especially as engine parameters evolve with mission profiles. To address this, our study evaluates the dynamic reliability of reusable rocket engines (RREs) and their subcomponents under different failure modes using multi-physics system-level modelling and simulation, with a particular focus on turbopump components. Transient condition modelling and performance analysis, conducted using EcosimPro-ESPSS software (version 6.4.34), revealed that turbopump components maintain high reliability under nominal conditions, with turbine blades demonstrating significant fatigue life even under varying thermal and mechanical loads. Additionally, the proposed predictive model estimates the remaining useful life of critical components, offering valuable insights for improving the longevity and reliability of turbopumps in reusable rocket engines. This study employs deterministic, thermally dependent structural simulations, with key control objectives including end-state tracking of combustion chamber pressure and mixture ratios and the verification of operational constraints, exemplified by the LUMEN demonstrator engine and the LE-5B-2 engine class.

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Simplício, Pedro, Andrés Marcos und Samir Bennani. „Reusable Launchers: Development of a Coupled Flight Mechanics, Guidance, and Control Benchmark“. Journal of Spacecraft and Rockets 57, Nr. 1 (Januar 2020): 74–89. http://dx.doi.org/10.2514/1.a34429.

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Bonnal, Ch, und M. Caporicci. „Future reusable launch vehicles in europe: the FLTP (Future Launchers Technologies Programme)“. Acta Astronautica 47, Nr. 2-9 (Juli 2000): 113–18. http://dx.doi.org/10.1016/s0094-5765(00)00050-3.

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Dissertationen zum Thema "Reusable Launchers"

Berry, W. „Reusable launchers“. Thesis, Cranfield University, 1993. http://hdl.handle.net/1826/3902.

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This research on Reusable Launchers was motivated by the need to reduce substantially the cost of space transportation. The specific objective was to- explore the perception that launcher reusability is the key to achieving these major cost reductions. The exploration was achieved by undertaking a comparative system study on potentially feasible reusable launcher concepts, using a consistent set of design tools, a standard analysis methodology and a standard reference mission. To set the background f or the research, the results of an extensive literature review are 'presented on the vehicle studies and technology developments that are engaged across the world on reusable launchers. Comprehensive vehicle studies appear to be engaged without justification for the choice of selected concepts in the absence of results from comparative system studies of reusable launchers. Technology developments also appear to be engaged without clear links to needs derived from vehicle system studies. The challenge of reusability is then addressed. Firstly, to set the performance and cost targets of reusable launchers, the capabilities of current expendable launchers are derived. Secondly, to establish the operational requirements for reusable launchers, the probable space transportation needs for the early 21st century are derived. Thirdly, the concepts and characteristics of reusable launchers are derived, allowing the selection, on a rationale basis, of a short-list of 13 potentially feasible reusable launcher concepts for analysis in the research. The performance equations of reusable launchers are 'then derived, leading to the preparation of the comparative analysis tools. The major work-of the research, which ''comprises the performance analysis, technical feasibility assessment and cost"analysis of each candidate vehicle are, then presented and compared-. A set of acceptance requirements for performance,, technical feasibility and operational costs - of reusable launchers is then -derived. The results of the comparative analysis for each candidate launcher are then measured , against these requirements. The results of the comparative analysis show that only 2 of the' 13 candidate reusable launcher concepts are able to meet all the acceptance'requirements. These two acceptable vehicles are both rocket-propelled. They are, ýin order of preference: a single-stage-to-orbit, rocket-propelled, vertical launch and vertical landing vehicle; a two-stage-to-orbit, rocket-propelled, vertical launch and horizontal landing vehicle. The operational ''costs per launch for these two'vehicles,, based on a utilisation plan of 3 vehicles operating for 20 years at a launch rate of 12 launches per year, was calculated to be about 20 % of the current costs of the European Ariane 44L expendable launcher. This warrants their further evaluation in a thorough feasibility study. The more complex, air-breathing propelled, horizontal launch and landing vehicles were found to be unable to meet the performance, technical feasibility and cost requirements: Several vehicles were found to be unable to deliver a positive payload mass to orbit; Several vehicles were found to have technology requirements that were deemed to be infeasible to achieve; Several vehicles were found to have operational costs ranging from equal to double that -of the European Ariane 44L expendable launcher,, which -was- adopted as a comparative reference vehicle. The contributions of this research to the advancement of knowledge on reusable launchers are: a clear identification of the performance, capability limits of 13 plausible reusable launcher concepts; an analysis methodology for determining the performance capability limits for any reusable launcher concept; a clear identification of the reasons. for the poor practical performance of air-breathing propulsion systems for Earth-to-orbit launchers, which results from their installed operational characteristics.

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Zaragoza, Prous Guillermo. „Guidance and Control for Launch and Vertical Descend of Reusable Launchers using Model Predictive Control and Convex Optimisation“. Thesis, Luleå tekniska universitet, Institutionen för system- och rymdteknik, 2020. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-81354.

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The increasing market of small and affordable space systems requires fast and reliablelaunch capabilities to cover the current and future demand. This project aims to studyand implement guidance and control schemes for vertical ascent and descent phases ofa reusable launcher. Specifically, the thesis focuses on developing and applying ModelPredictive Control (MPC) and optimisation techniques to several kino-dynamic modelsof rockets. Moreover, the classical MPC method has been modified to include a decreasingfactor for the horizon used, enhancing the performance of the guidance and control.Multiple scenarios of vertical launch, landing and full fligth guidance on Earth and Mars,along with Monte Carlo analysis, were carried out to demonstrate the robustness of thealgorithm against different initial conditions. The results were promising and invite tofurther research.

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Gibart, Jules. „Non-linear stability of a liquid propelled rocket engine in closed loop regulation“. Electronic Thesis or Diss., université Paris-Saclay, 2024. http://www.theses.fr/2024UPAST110.

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Dans le cadre du développement de moteurs de fusée réutilisables, les exigences de fonctionnement des différents éléments composant un moteur ont connu de grandes évolutions. Alors qu'un moteur classique était conçu pour un nombre restreint de points de fonctionnement, un moteur réutilisable doit répondre à des exigences sur une large plage de points, afin d'effectuer des manoeuvres plus complexes. En conséquence, les lois de commande des moteurs fusées ont subi une évolution similaire, rendant nécessaire la loi de commande en boucle fermée. Bien que de nombreuses études aient été réalisées sur des lois de commande, peu de travaux portent sur la stabilité du moteur en boucle fermée. Dans cette optique, l'objectif de ces travaux est de proposer une démonstration de stabilité d'un modèle de moteur fusée, ainsi qu'un contrôleur permettant d'obtenir des garanties de stabilité du modèle. En premier lieu, un modèle typique de moteur de fusée à ergols liquide est développé, sous forme d'espace d'états. Ce type de modèle, bien que plus courant, se révèle peu adapté à l'étude de la stabilité, de par sa formulation hautement non-linéaire. Dans ce cadre, l'utilisation d'une fonction de Lyapunov se révèle complexe, et une reformulation du modèle est envisagée, sous forme d'un modèle Hamiltonien à ports. Un second chapitre permet d'introduire la notion de modèle Hamiltonien à ports. Ce type de modèle met en valeur les transferts énergétiques qui ont lieu entre les différents éléments d'un système, et sont construits avec une structure géométrique fixe. Ces différentes caractéristiques permettent une étude directe de la passivité d'un système, un outil d'analyse de la stabilité d'un système. La reformulation permet de trouver une fonction caractéristique d'un système Hamiltonien à ports, l'Hamiltonien, qui prouve la passivité d'un système et peut être formulé comme une fonction de Lyapunov. Cette démonstration donne des conditions de stabilité sur la modélisation du système, ainsi que sur le contrôleur appliqué en boucle fermée. Dans le cas où la démonstration directe de passivité n'est pas réalisable, un contrôleur peut être construit pour assurer la passivité de la boucle fermée. Pour conférer les propriétés de la passivité au modèle de moteur utilisé, la théorie du contrôle par passivité est présentée. Le principe d'un tel contrôleur est d'assurer la stabilité d'un système en rendant la boucle fermée passive. Avec la théorie des systèmes Hamiltonien à ports cependant, ce contrôleur permet aussi de modifier la structure géométrique hamiltonienne, afin de reformuler un système sous forme Hamiltonienne à ports. Ce contrôleur permet de rendre le système passif autour d'un point de fonctionnement désiré par l'utilisateur, qui peut être changé au cours du temps. Ainsi, ce contrôleur permet un suivi de trajectoire avec des garanties de passivité du système au cours du temps. Le quatrième chapitre propose une approche différente pour établir un contrôleur stabilisant, à l'aide de la théorie de la contraction. La propriété de contraction d'un système dénote sa capacité à converger rapidement vers une trajectoire de référence. Cette propriété constitue une forme de stabilité exponentielle, plus puissante que la stabilité par passivation. Le contrôleur peut de plus être réalisé aisément, en résolvant des inégalités linéaires matricielles. Enfin, les résultats de ces travaux sont présentés à l'aide de simulations sur MATLAB Simulink, et permettent de conclure sur les différents contrôleurs présentés. Un contrôleur simple proportionnel intégral dérivé (PID) est construit pour permettre une comparaison. Les résultats montrent que les contrôleurs réalisés proposent des propriétés stabilisantes, alors que le contrôleur PID est instable dans certaines zones de fonctionnement. Le contrôleur par passivité étend le domaine de stabilité du système, et le contrôleur par contraction empêche le système de quitter le domaine de stabilité du système original
With the development of reusable rocket engines, the operating requirements of the various components in an engine have significantly increased. While a non-reusable engine was designed for a limited number of operating points, a reusable engine must meet requirements over a wide range of points to perform complex maneuvers. Consequently, rocket engine control laws have evolved similarly, with the introduction of closed-loop control laws. Although many studies have been conducted on control laws, few works focus on the stability of the engine in closed-loop control. In this context, the objective of this work is to propose a demonstration of the stability of a rocket engine model, as well as a controller that guarantees the stability of the model. First, a model of a liquid propelled rocket engine is proposed under a state-space form. Although more common, this type of modeling does not allow for an easy stability analysis due to its highly nonlinear terms. In this context, the use of a Lyapunov function proves to be cumbersome, and a reformulation of the model is considered, in the form of a Port-Hamiltonian model, more suited for stability analysis of the system. A second chapter introduces the concept of the Port-Hamiltonian model. This type of model highlights the energy transfers that occur between the various components of a system and is built with a fixed geometric structure. These characteristics allow for a direct study of the passivity of a system, a tool for stability analysis the stability. The reformulation allows for the identification of a characteristic function of a Port-Hamiltonian system, the Hamiltonian function, which can be used to prove the passivity of a system and can be formulated as a Lyapunov function. This demonstration provides stability conditions for the system as well as the controller applied in the closed-loop system. In cases where a direct demonstration of passivity is not possible, a controller can be constructed to ensure the passivity of the closed-loop system. To endow the rocket engine model with passivity properties, the third chapter presents passivity-based control (PBC) theory. The principle of such a controller is to ensure the stability of a system by making the closed-loop system passive. Coupled with Port-Hamiltonian systems theory, however, this controller also allows for modification of the Hamiltonian geometric structure to reformulate a system into Port-Hamiltonian form. This controller makes the system passive around a desired operating point, which can be changed over time. Thus, this controller enables trajectory tracking with passivity guarantees over time. The fourth chapter proposes a different approach to establish a stabilizing controller using contraction theory. The contraction property of a system indicates its ability to rapidly converge towards a reference trajectory. This property represents a form of exponential stability, which is more robust than stability through passivation. Moreover, the controller can be easily implemented by solving linear matrix inequalities. Finally, the results of this work are presented through simulations on MATLAB Simulink, allowing for conclusions on the various controllers presented. A simple proportional-integralderivative (PID) controller is constructed for comparison. The results show that the designed controllers offer stabilizing properties, while the PID controller is unstable in certain operating regions. The passivity-based controller extends the stability domain of the system, and the contraction-based controller prevents the system from leaving the stability domain of the original system

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MABBOUX, ROMAIN. „Optimization of the pressurization system of the Themis reusable rocket first stage demonstrator“. Thesis, KTH, Skolan för industriell teknik och management (ITM), 2020. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-301295.

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The design of complex systems such as a launcher, or subsystems like its pressurization system, is known to be fastidious and expensive especially in the space domain. However, the recent emergence of new actors in this domain has been a game changer, binding these systems to fulfil more and more requirements (cheap, efficient, rapidly produced, aesthetic, environmentally friendly…) in order to compete with the market. For these reasons, engineers now need more than ever to consider the full picture of a system or subsystem in order to optimize it. In this context, this document aims to present the method and the results obtained in the optimization of the pressurization system of a reusable rocket first stage named Themis 3. The modelling of the pressurization system has been realized through a software called Geeglee. In order to cover as many impacts of this system as possible, an important part of the rocket stage has been considered and modelled (from the pressurization gases to the propellant tanks, passing by the pressurant and propellant feeding lines). Three High Level Requirements (HLR) have been identified as of major importance for the trade-off in the design of the pressurization system: the total mass impact, the total Recursive Cost (RC) impact and the total Non-Recursive Cost (NRC) impact. This optimization has in particular permitted to confirm some well-known results, so to say that exogenous pressurization systems represent a smaller mass impact on the vehicle, at the expense of a higher RC compared with autogenous systems.
Utformningen av komplexa system, som till exempel en bärraket, eller delsystem som dess trycksättningssystem, är känd för att vara krävande och kostsam, särskilt inom rymdteknikområdet. Den senaste tidens uppkomst av nya aktörer på detta område har dock förändrat spelregler då dessa system tvingas uppfylla allt fler krav (billiga, effektiva, snabbt producerade, estetiska, miljövänliga etc.) kunna vara konkurrenskraftiga på marknaden. Av dessa skäl måste ingenjörer nu mer än någonsin beakta hela bilden av ett system eller delsystem för att kunna optimera det. I detta sammanhang syftar detta dokument till att presentera metoden och resultaten från optimeringen av trycksättningssystemet för en återanvändbar rakets första steg som heter Themis 3. Modelleringen av trycksättningssystemet har genomförts med hjälp av ett systemmodelleringsverktyg som kallas Geeglee. För att täcka in så många effekter av detta system som möjligt har en viktig del av raketsteget beaktats och modellerats (från trycksättningsgaserna till drivmedelstankarna, via tryck- och drivmedelsmatningsledningarna). Tre krav på hög nivå har identifierats som mycket viktiga för avvägningen vid utformningen av trycksättningssystem konstruktion: den totala masspåverkan, den totala icke-rekursiva kostnaden och den totala rekursiva kostnaden. Denna optimering har framför allt gjort det möjligt att bekräfta vissa välkända resultat, dvs. att exogena trycksättningssystem ger en mindre masspåverkan på fordonet, på bekostnad av en högre RC jämfört med autogena system.

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Bulut, Jane. „Design and CFD analysis of the demonstrator aerospike engine for a small satellite launcher application“. Master's thesis, Alma Mater Studiorum - Università di Bologna, 2020.

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Starting with a brief overview of thrust generation for launchers, this study focuses on the design process of the demonstrator aerospike engine, DEMOP-1, of the Pangea Aerospace's commercial grade engine and its flow field analysis. The primary goal of the study is to obtain the plug nozzle design delivers 30 kN thrust using cryogenic liquid oxygen (LOX) as the oxidizer and cryogenic liquid methane (LCH4) as the fuel, with the mixture ratio of 3.4. Design parameters considered as 30 bar of combustion chamber pressure (Po) and expansion ratio as 15 for an optimum expanded nozzle. On the basis of decided design characteristics, Angelino's method is used to design the nozzle contour through MATLAB. The flow field over the aerospike analyzed using commercial CFD program FLUENT for sea level, optimum expansion and vacuum conditions. Flow simulations are carried out for air (specific heat ratio, gamma= 1.4), and afterwards based on the obtained thrust values at each altitude for air, expected thrust values for the real propellant, LOX/LCH4 (specific heat ratio, gamma = 1.1664), are calculated. Finally, the study is concluded with the comparison of trend in thrust and specific impulse for conventional bell nozzle and aerospike. For the conventional bell engine the values obtained in commercial computational simulation of chemical rocket propulsion and combustion software RPA for bell nozzle with same characteristics with aerospike, Po = 30 bar and expansion ratio = 15, are taken as reference for sea level, optimum expansion level and vacuum condition performance. Due to its ability to adopt the altitude, aerospike delivers higher performance at the low altitudes with respect to the conventional bell nozzle which has the same expansion ratio and combustion chamber pressure. Last in order but not in importance, after obtaining the flow field on plug of the aerospike, the shock wave impingement on the nozzle surface at sea level has been investigated.

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Lantelme, Melissa. „Modélisation des grandeurs aérothermodynamiques pariétales : application à la rentrée atmosphérique des lanceurs réutilisables“. Electronic Thesis or Diss., Toulouse, ISAE, 2024. http://www.theses.fr/2024ESAE0025.

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Cette étude consiste à développer une méthode de prévision de la distribution du flux de chaleur surfacique pour des lanceurs réutilisables qui soit adaptée à une phase de conception préliminaire. Les phases de vol hypersonique les plus critiques lors de la réentrée atmosphérique du lanceur sont celles pour lesquelles les flux de chaleurs les plus importants apparaissent à la surface. Ainsi, pour les études de conception et de préparation de mission, il est indispensable de posséder des outils élaborés et rapides qui permettent d’obtenir des estimations de ces flux de chaleur surfaciques. Cela permet ainsi un dimensionnement rapide du design étudié car le flux de chaleur est une données d’entrée pour une multitude de spécialités: l’aérothermodynamique, l’optimisation de la trajectoire, la conception structurelle, etc. L’objectif principal de cette thèse est d’évaluer si l’utilisation des méthodes de « Machine Learning » apporte une plus-value pour le développent d’un modèle de prévision de la distribution du flux de chaleur convectif-diffusif surfacique. Pour cela, nous proposons une méthode qui consiste à développer deux modèles de substitution consécutifs. Le premier modèle de substitution est établi pour effectuer un adimensionnement des variables. Cela nous permet par la suite d’utiliser le modèle indépendamment des conditions de vol et d’altitude. Le deuxième modèle de substitution basé sur les réseaux de neurones établi la corrélation entre les variables d’entrée – topologie de pression surfacique et variables géométriques – et le flux de chaleur en sortie. Dans le cadre de cette thèse la méthode a été testée et développée pour un écoulement hypersonique laminaire en régime continue. Pour analyser ses capacités et limites, nous avons appliqué cette méthode sur l’étage orbital duconcept de lanceur « SpaceLiner ». Une base de données est établie avec les calculs de Navier-Stokes sur un ensemble de points de vols et d’orientations choisi. Les variables d’entrées sont adimensionnées à partir d’équations existantes ou de transformations mathématiques simples. Pour l’adimensionnement de flux de chaleur un modèle de substitution pour la prévision du flux de chaleur au point d’arrêt pour notre domaine d’intérêt est développé. Ce modèle est basé sur le modèle de Lepage-Vérant et du Krigeage. Cela nous permet d’obtenir une prévision avec une erreur moyenne relative de 1.9% par rapport aux résultats de CFD. Le développement puis l’évaluation du modèle de substitution pour la prévision du flux de chaleur en tout point de la paroi sont effectués. Nous avons comparé nos résultats à ceux issus de calculs CFD à la fois dans des zones convexes et –partiellement– dans des zones planes de façon satisfaisante avec une précision supérieure ou égale à celle des méthodes existantes. Des réseaux de neurones entrainés sur un petit nombre de points de vol sont capables d’interpoler et d’extrapoler les résultats à d’autres points de vol à proximité. Cependant, dans des zones à flux de chaleur adimensionné très faible (inférieur à 0.05) le modèle reste peu fiable mais cet aspect n’est pas essentiel dans une étude de conception préliminaire. De plus, la base de données n’est pas adaptée au développement d’un modèle de substitution pour la zone concave. Comme les modèles existants, la méthode proposée ne réussit pas à prédire les phénomènes physiques complexes comme des interactions choc-choc. En conclusion la méthode proposée a démontré son potentiel d’intégration dans des phases de conception préliminaire
For reusable launch vehicles critical aerothermal loads occur on the descent trajectory in the hypersonic continuum regime. Thus, for the pre-design phase of the mission preparation, it is essential to have advanced models with low response time (CPU time) at one’s disposal to predict the heat fluxes on the vehicle’s surface. This is necessary to make informed decisions on trade-offs required between system, aerothermal and trajectory optimisation aspects. The objective of this thesis is to investigate whether machine learning techniques offer additional value in the development of such a model predicting the wall heat flux distribution. For this, we propose a method which consists of two consecutive surrogate models. The first one provides a method of nondimensionalisation, to enable the prediction of heat flux independent of the flight points at different altitudes and freestream conditions. The second surrogate model maps four geometric and pressure based dimensionless input variables to the dimensionless heat flux output variable. This correlation is modelled with neural networks. In order to analyse the prediction accuracy, capabilities and limitations, we apply the proposed method to the SpaceLiner concept. For this, we build a database with Computational fluid dynamics (CFD) simulations which contains data from different flight points and orientations of the SpaceLiner orbiter stage. The nondimensionalisation model of the stagnation point heat flux is build based on the Lepage-Vérant model and Kriging and permits its prediction with an accuracy of a mean relative error of 1.9% compared to the results of CFD equations. The surrogate model based on neural networks for the prediction of the heat flux distribution at the current development state complies to our objective for the convex zones and partially for the flat zones if we disregard the very low heat flux zones and complex physical phenomena such as shock-shock interactions. These results ensure better or equivalent prediction accuracy as existing pre-design tools. The database was unsuited to developed a reliable model for the concave zones. Overall interpolation and close proximity extrapolation to different flight points for the given vehicle is possible. However further generalisation towards different vehicle shapes remain unreachable with the current setup. The proposed method has demonstrated potential to be integrated in the pre-design process

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Bücher zum Thema "Reusable Launchers"

Gibbins, Martin N. Systems integration and demonstration of advanced reusable structure for ALS. Hampton, Va: National Aeronautics and Space Administration, Langley Research Center, 1991.

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Center, NASA Glenn Research, Hrsg. Advanced electric propulsion for RLV launched geosynchronous spacecraft. [Cleveland, Ohio]: National Aeronautics and Space Administration, Glenn Research Center, 1999.

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Advanced electric propulsion for RLV launched geosynchronous spacecraft. [Cleveland, Ohio]: National Aeronautics and Space Administration, Glenn Research Center, 1999.

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Reentry: SpaceX, Elon Musk, and the Reusable Rockets that Launched a Second Space Age. BenBella Books, Inc., 2024.

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Buchteile zum Thema "Reusable Launchers"

van Pelt, Michel. „Reusable launchers“. In Dream Missions, 45–85. Cham: Springer International Publishing, 2017. http://dx.doi.org/10.1007/978-3-319-53941-6_3.

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Salt, David J. „Could a Subsonic Air-Launched Reusable Launch Vehicle (RLV) Enable a Paradigm Shift in Space Operations?“ In Space Operations: Innovations, Inventions, and Discoveries, 185–217. Reston, VA: American Institute of Aeronautics and Astronautics, Inc., 2015. http://dx.doi.org/10.2514/5.9781624101991.0185.0218.

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Kislykh, V. V., und I. A. Reshetin. „Afterbody Effects Study on Energia Reusable Launcher Models. Selection of Jet Propulsive Masses Parameters Within Jet Streams Ejecting Model“. In Separated Flows and Jets, 851–54. Berlin, Heidelberg: Springer Berlin Heidelberg, 1991. http://dx.doi.org/10.1007/978-3-642-84447-8_105.

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Yeager, Kenneth R. „Program Evaluation: This Is Rocket Science“. In Evidence-Based Practice Manual: Research and Outcome Measures in Health and Human Services, 647–53. Oxford University PressNew York, NY, 2004. http://dx.doi.org/10.1093/oso/9780195165005.003.0071.

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Abstract How could the space shuttle Challenger catastrophe and the deaths of all crew members have been prevented? How safe is health care? Can medical errors be prevented? Can process measurement and effective quality programming prevent unnecessary fatalities? This chapter examines processes associated with estimation of safety in arenas where minimal tolerance exists for error. Process measurement and evaluation is recommended as a tool to contribute to reduction of critical errors. The space shuttle program was begun in the early 1970s with the concept of creating reusable craft for transporting people and cargo into space. When the first shuttle, Columbia, was launched in 1981, it represented the realization of a new era of reusable spacecraft. One year following the introduction of Columbia, the space shuttle Challenger rolled off the assembly line as the second of the new U.S. fleet of reusable spacecraft. Two others were to follow: Discovery in 1983 and Atlantis in 1985.

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Koshova, Svitlana. „Innovative Trends in the Creation of Rocket and Spacecraft Equipment for the Purpose of Enhancement of National Security“. In Стратегія сучасного розвитку України: синтез правових, освітніх та економічних механізмів : колективна монографія / за загальною редакцією професора Старченка Г. В., 215–27. ГО «Науково-освітній інноваційний центр суспільних трансформацій», 2022. http://dx.doi.org/10.54929/monograph-12-2022-05-01.

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The space industry has grown and evolved significantly over the past six decades. Exploitation of space under the auspices of several space states has become an activity in more than 60 countries around the world. For its financing, in addition to state capital, the private sector began to be actively involved in the form of investment resources, which made it possible to increase the pace of its development several times. In recent years, interest in space exploration and the benefits that can be, derived from it has grown significantly. This became possible due to the significant simplification of travel beyond the borders of the Earth under the influence of innovative trends. Therefore, the current trends in the global space industry are, defined in the scientific study. The role of the space sphere in the general development of the country and the formation of its competitive advantages has been, established. The key directions of space travel in our time are outlined, and the role of reusable launch systems for orbital vehicles in their actualization. Described the SN20 reusable launch system developed by SpaceX as a basis for creating opportunities for travel to Mars. Its technical characteristics and the sequence of its licensing procedure for the first launch are, defined. The dynamics of changes in the cost of rocket launches under the influence of innovations were, studied. The reasons for the resumption of trips to the Moon are established, a parallel is, drawn between, the influence, of such trips on the further colonization of Mars. The current trends in the field of satellite launches have been determined, the number of orbital launch attempts by countries has been, analyzed, and the forecast for changes in the number of satellites until 2030 has been determined. The problem of combating space debris, which remained from past missions and will remain from future missions, was, studied. Directions for using space technologies to fight global warming are, defined. Ways of using 3D printing for the needs of future space missions are, analyzed. The structure of the direction of venture capital in space technologies in terms of the countries, as well as in terms of the stages of implementation of space projects in terms of value and quantity, were, studied.

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Konferenzberichte zum Thema "Reusable Launchers"

Denaro, Angelo, Elena Brach Prever und Marco Nebiolo. „Developments on Cryogenic Insulations for Reusable Launchers“. In AIAA/CIRA 13th International Space Planes and Hypersonics Systems and Technologies Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2005. http://dx.doi.org/10.2514/6.2005-3436.

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Denaro, A., E. Brach Prever, M. Nebiolo und H. Ritter. „Screening Tests on Cryogenic Insulations for Reusable Launchers“. In International Conference On Environmental Systems. 400 Commonwealth Drive, Warrendale, PA, United States: SAE International, 2005. http://dx.doi.org/10.4271/2005-01-2899.

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Denaro, A., und H. Ritter. „Developments on Cryogenic Tank Insulation for Reusable Launchers“. In International Conference On Environmental Systems. 400 Commonwealth Drive, Warrendale, PA, United States: SAE International, 2004. http://dx.doi.org/10.4271/2004-01-2565.

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Liu, Zibo, und Ran Zhang. „Actuator-constrained Trajectory Optimization for Reusable Launchers’ Landing“. In 2022 13th Asian Control Conference (ASCC). IEEE, 2022. http://dx.doi.org/10.23919/ascc56756.2022.9828171.

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Angelino, E., D. Dosio und G. Borriello. „Reusable launchers tank structures - Requirements definition and experimental development“. In 8th AIAA International Space Planes and Hypersonic Systems and Technologies Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1998. http://dx.doi.org/10.2514/6.1998-1592.

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Wulz, H., U. Trabandt, H. Wulz und U. Trabandt. „Large integral hot CMC structures designed for future reusable launchers“. In 32nd Thermophysics Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1997. http://dx.doi.org/10.2514/6.1997-2485.

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Sippel, Martin, Arnin Herbertz und Holger Burkhardt. „Reusable Booster Stages: A Potential Concept for Future European Launchers“. In AIAA/CIRA 13th International Space Planes and Hypersonics Systems and Technologies Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2005. http://dx.doi.org/10.2514/6.2005-3242.

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Kostromin, S. „Cost effectiveness estimates of the partially reusable launchers family with uniform components“. In 9th International Space Planes and Hypersonic Systems and Technologies Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1999. http://dx.doi.org/10.2514/6.1999-4887.

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Sudmeijer, Kees, Arjen Kloosterman, Benedict Lefeber und Cyril Wentzel. „Technology Development for Metallic Hot Structures in Aerodynamic Control Surfaces of Reusable Launchers“. In AIAA/AAAF 11th International Space Planes and Hypersonic Systems and Technologies Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2002. http://dx.doi.org/10.2514/6.2002-5161.

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Iannelli, Andrea, Dimitris Gkouletsos und Roy S. Smith. „Robust Control Design for Flexible Guidance of the Aerodynamic Descent of Reusable Launchers“. In AIAA SCITECH 2023 Forum. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2023. http://dx.doi.org/10.2514/6.2023-2171.

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