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Littérature scientifique sur le sujet « Upper stage space rocket engine »

Auteur : Grafiati

Publié le 4 juin 2021

Mis à jour le 1 février 2022

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Articles de revues sur le sujet "Upper stage space rocket engine"

Timoshenko, V. I., V. P. Halynskyi et Yu V. Knyshenko. « Theoretical studies on rocket/space hardware aerogas dynamics ». Technical mechanics 2021, n^o 2 (29 juin 2021) : 46–59. http://dx.doi.org/10.15407/itm2021.02.046.

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This paper presents the results of theoretical studies on rocket/space hardware aerogas dynamics obtained from 2016 to 2020 at the Department of Aerogas Dynamics and Technical Systems Dynamics of the Institute of Technical Mechanics of the National Academy of Sciences of Ukraine and the State Space Agency of Ukraine along the following lines: rocket aerodynamics, mathematical simulation of the aerogas thermodynamics of a supersonic ramjet vehicle, jet flows, and the hydraulic gas dynamics of low-thrust control jet engines. As to rocket aerodynamics, computational methods and programs (CMPs) were developed to calculate supersonic flow past finned rockets. The chief advantage of the CMPs developed is computational promptness and ease of adding wings and control and stabilization elements to rocket configurations. A mathematical simulation of the aerogas thermodynamics of a supersonic ramjet vehicle yielded new results, which made it possible to develop a prompt technique for a comprehensive calculation of ramjet duct flows and generalize it to 3D flow past a ramjet vehicle. Based on marching methods, CMPs were developed to simulate ramjet duct flows with account for flow past the airframe upstream of the air inlet, the effect of the combustion product jet on the airframe tail part, and its interaction with a disturbed incident flow. The CMPs developed were recommended for use at the preliminary stage of ramjet component shape selection. For jet flows, CMPs were developed for the marching calculation of turbulent jets of rocket engine combustion products with water injection into the jet body. This made it possible to elucidate the basic mechanisms of the effect of water injection, jet–air mixing, and high-temperature rocket engine jet afterburning in atmospheric oxygen on the flow pattern and the thermogas dynamic and thermalphysic jet parameters. CMPs were developed to simulate the operation of liquid-propellant low-thrust engine systems. They were used in supporting the development and ground firing tryout of Yuzhnoye State Design Office’s radically new system of control jet engines fed from the sustainer engine pipelines of the Cyclone-4M launch vehicle upper stage. The computed results made it possible to increase the informativity of firing test data in flight simulation. The CMPs developed were transferred to Yuzhnoye State Design Office for use in design calculations.

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Simmons, J., et Richard Branam. « Parametric Study of Dual-Expander Aerospike Nozzle Upper-Stage Rocket Engine ». Journal of Spacecraft and Rockets 48, n^o 2 (mars 2011) : 355–67. http://dx.doi.org/10.2514/1.51534.

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Schmehl, Roland, et Johan Steelant. « Computational Analysis of the Oxidizer Preflow in an Upper-Stage Rocket Engine ». Journal of Propulsion and Power 25, n^o 3 (mai 2009) : 771–82. http://dx.doi.org/10.2514/1.38309.

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Beyer, Steffen, Stephan Schmidt, Franz Maidl, Rolf Meistring, Marc Bouchez et Patrick Peres. « Advanced Composite Materials for Current and Future Propulsion and Industrial Applications ». Advances in Science and Technology 50 (octobre 2006) : 174–81. http://dx.doi.org/10.4028/www.scientific.net/ast.50.174.

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Various technology programmes in Europe are concerned with preparing for future propulsion technologies to reduce the costs and increase the life time of components for liquid rocket engine components. One of the key roles to fulfil the future requirements and for realizing reusable and robust engine components is the use of modern and innovative materials. One of the key technologies which concern various engine manufacturers worldwide is the development of fibrereinforced ceramics – CMC's (Ceramic Matrix Composites). The advantages for the developers are obvious – the low specific weight, the high specific strength over a large temperature range, and their good damage tolerance compared to monolithic ceramics make this material class extremely interesting as a construction material. Different kind of composite materials are available and produced by EADS ST, the standard material SICARBON® (C/SiC made by Liquid Polymer Infiltration) and the new developed and qualified composite materials SICTEX® (C/SiC made by Liquid Silicon Infiltration) and CARBOTEX® (C/C made by Rapid Chemical Vapour Infiltration). The composites are based on textile techniques like weaving, braiding, stiching and sewing to produce multiaxial preforms, the SICTEX® material is densificated by the cost effective Liquid Silicon Infiltration (LSI). Over the past years, EADS Space Transportation (formerly DASA) has, together with various partners, worked intensively on developing components for airbreathing and liquid rocket engines. Since this, various prototype developments and hot firing-tests with nozzle extensions for upper and core stage engines and combustion chambers of satellite engines were conducted. MBDA France and EADS-ST have been working on the development of fuel-cooled composite structures like combustion chambers and nozzle extensions for future propulsion applications.

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Vinod, G., S. Renjith et V. Thaddeus Basker. « Thermo Structural Analysis of Carbon-Carbon Nozzle Exit Cone for Rocket Cryo Engines ». Applied Mechanics and Materials 877 (février 2018) : 320–26. http://dx.doi.org/10.4028/www.scientific.net/amm.877.320.

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Launch and space vehicle structures are required to be extremely weight efficient. The need to achieve the performance required for the engine in the upper stage of a launch vehicle, increase the payload capacity drives rocket engine manufacturers to seek higher thrust level, specific impulse and thrust to weight ratio. The use of high temperature C-C composite materials is an efficient way to reach these objectives by allowing use of high expansion ratio. Nozzle extensions benefiting of the outstanding thermal, mechanical and fatigue resistance of these materials to decrease mass and featuring high temperature margins. A three-directionally reinforced (3D) carbon-carbon (c-c) material nozzle exit cone is selected for the current study. C-C composite exit nozzle must possess excellent stability and strength under extreme conditions for a specified amount of time. Carbon-carbon composites are appropriate materials for applications that require high specific strength at elevated temperatures. The paper describes the thermo structural analysis of a typical c/c nozzle exit cone.

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Rugescu, Radu D., Dragos Ronald Rugescu et Efim Micu. « Orbital Launcher NERVA as the First Proof of the Discontinuous Variational Solution for the Atmospheric Ascent ». Applied Mechanics and Materials 555 (juin 2014) : 91–101. http://dx.doi.org/10.4028/www.scientific.net/amm.555.91.

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Since the earliest days of astronautics, more than a century ago, low cost space launchers persevered to be a long desire for the space flight thinkers. Once space flight became a daily business along the late `50-s, first by consuming large financial resources, the interest for cheap space launchers became even more laud. Today’s growing interest in small satellites have bolstered a large series of space technology companies including Virgin Galactic Corp., Garvey Spacecraft Corp., Quantum Research International, Ventions LLC, Sierra Nevada Corp., Generation Orbit Launch Services and even the giant Boeing to work on the development of various types of such vehicles, some of them of actually small size. They have announced recent progresses in their efforts to develop and test small-satellite launchers and rocket engines. Romanian space launcher effort includes the NERVA project, with the ORVEAL compound engine for the upper stage, securing the orbital injection, project developed by the team of professors and researchers from ADDA Ltd, Bucharest. This project is based on a series of innovative concepts, including the optimal ascent program first proposed by the ADDA team by means of the new discontinuous variational optimization, which is here described in detail.

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Дронь, Николай Михайлович, Кирилл Валерьевич Коростюк, Александр Вячеславович Голубек, Людмила Григорьевна Дубовик et Алексей Владимирович Кулик. « ОЦЕНКА ВОЗМОЖНОСТЕЙ ПРИМЕНЕНИЯ СУБОРБИТАЛЬНЫХ РАКЕТ-НОСИТЕЛЕЙ ДЛЯ ВЫВЕДЕНИЯ СРЕДСТВ УВОДА ОБЪЕКТОВ КОСМИЧЕСКОГО МУСОРА С НИЗКИХ ОКОЛОЗЕМНЫХ ОРБИТ ». Aerospace technic and technology, n^o 4 (28 août 2020) : 60–65. http://dx.doi.org/10.32620/aktt.2020.4.07.

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The article is devoted to an actual problem of clearing of low earth orbits from space objects of a technogenic origin. Existing versions of struggle against space debris, in particular, removal of technogenic objects with help of the special means for deorbiting delivered into a target orbit by launch vehicles that are especially actual for bulky space debris are considered. Recognizing that the ascent of such means for deorbiting by orbital launch vehicles demands large financial expenses, for an increase of efficiency of delivery the means for deorbiting on a low earth orbit widely known sub-orbital launch vehicles are offered: MAXUS, TEXUS (VSB-30), REXUS (Improved Orion), SS-520, MH-300, Black Brant 12А and the estimation of a capability of their application also is conducted. Are considered the use of sub-orbital launch vehicles for the ascent the means for deorbiting on altitudes of a concentration of space debris on a low earth orbit on a trajectory, close to vertical, with the subsequent operations of interception of demanded space objects, and also modernization of launch vehicles by addition of an additional stage. Results of calculations of an injection trajectory of the means for deorbiting in weight in a layer of space debris in altitude 600 … 1200 km showed of 150 kg that sub-orbital launch vehicles MAXUS, SS-520, Black Brant 12A allow executing delivery the means for deorbiting to altitudes from 770 km to 1200 km and to supply time of its presence in a layer of space debris 420 … 850 s. The most perspective sub-orbital rocket is MAXUS. It possesses higher power and a capability of installation of an additional stage by a decrease in weight of a payload with small losses the power of the first stage. It is shown that the given configuration of the rocket with engine thrust specific impulse in vacuum 300 s and engine thrust in vacuum 16 кН is capable to inject into an elliptical orbit with an altitude of apogee 600 km and altitude of a perigee 130 km with a corner of an inclination 5,5 degrees payload in weight of 55 kg. For orbit short circuit in apogee at the altitude, the upper stage should supply 600 km increase the speeds, equal 133 m/s. Mass characteristics of the second stage are induced.

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Alexander O., Mayboroda. « Space Debris Removal and Exploitation of Lunar Resources - Profitability Perspectives ». AEROSPACE SPHERE JOURNAL, n^o 2 (26 juin 2021) : 24–33. http://dx.doi.org/10.30981/2587-7992-2020-107-2-24-33.

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Removing at least half of space debris, consisting of large metal objects, could be a cost-effective activity if, first, it becomes raw material, along with other extraterrestrial resources such as regolith, for the production of heat shields which is aimed to return the upper stages of medium and heavy rockets and, secondly, as a working medium for low-thrust electric rocket engines of interorbital tugs. Heat shields from external resources provide a means of increasing the payload of reusable launch vehicles.

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Rubio Hervas, Jaime, et Mahmut Reyhanoglu. « Thrust Vector Control of an Upper-Stage Rocket with Multiple Propellant Slosh Modes ». Mathematical Problems in Engineering 2012 (2012) : 1–18. http://dx.doi.org/10.1155/2012/848741.

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The thrust vector control problem for an upper-stage rocket with propellant slosh dynamics is considered. The control inputs are defined by the gimbal deflection angle of a main engine and a pitching moment about the center of mass of the spacecraft. The rocket acceleration due to the main engine thrust is assumed to be large enough so that surface tension forces do not significantly affect the propellant motion during main engine burns. A multi-mass-spring model of the sloshing fuel is introduced to represent the prominent sloshing modes. A nonlinear feedback controller is designed to control the translational velocity vector and the attitude of the spacecraft, while suppressing the sloshing modes. The effectiveness of the controller is illustrated through a simulation example.

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Casalino, Lorenzo, Filippo Masseni et Dario Pastrone. « Viability of an Electrically Driven Pump-Fed Hybrid Rocket for Small Launcher Upper Stages ». Aerospace 6, n^o 3 (14 mars 2019) : 36. http://dx.doi.org/10.3390/aerospace6030036.

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An electrically driven pump-fed cycle for a hybrid rocket engine is proposed and compared to a simpler gas-pressurized feed system. A liquid-oxygen/paraffin-based fuel hybrid rocket engine which powers the third stage of a Vega-like launcher is considered. Third-stage ignition conditions are assigned, and engine design and payload mass are defined by a proper set of parameters. Uncertainties in the classical regression rate correlation coefficients are taken into account and robust design optimization is carried out with an approach based on an epsilon-constrained evolutionary algorithm. A mission-specific objective function, which takes into account both the payload mass and the ability of the rocket to reach the required final orbit despite uncertainties, is determined by an indirect trajectory optimization approach. The target orbit is a 700 km altitude polar orbit. Results show that electrically driven pump-fed cycle is a viable option for the replacement of the conventional gas-pressurized feed system. Robustness in the design is granted and a remarkable payload gain is achieved, using both present and advanced technologies for electrical systems.

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Thèses sur le sujet "Upper stage space rocket engine"

Wahlström, Dennis. « Probabilistic Multidisciplinary Design Optimization on a high-pressure sandwich wall in a rocket engine application ». Thesis, Umeå universitet, Institutionen för fysik, 2017. http://urn.kb.se/resolve?urn=urn:nbn:se:umu:diva-138480.

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A need to find better achievement has always been required in the space industrythrough time. Advanced technologies are provided to accomplish goals for humanityfor space explorer and space missions, to apprehend answers and widen knowledges. These are the goals of improvement, and in this thesis, is to strive and demandto understand and improve the mass of a space nozzle, utilized in an upperstage of space mission, with an expander cycle engine. The study is carried out by creating design of experiment using Latin HypercubeSampling (LHS) with a consideration to number of design and simulation expense.A surrogate model based optimization with Multidisciplinary Design Optimization(MDO) method for two different approaches, Analytical Target Cascading (ATC) and Multidisciplinary Feasible (MDF) are used for comparison and emend the conclusion. In the optimization, three different limitations are being investigated, designspace limit, industrial limit and industrial limit with tolerance. Optimized results have shown an incompatibility between two optimization approaches, ATC and MDF which are expected to be similar, but for the two limitations, design space limit and industrial limit appear to be less agreeable. The ATC formalist in this case dictates by the main objective, where the children/subproblems only focus to find a solution that satisfies the main objective and its constraint. For the MDF, the main objective function is described as a single function and solved subject to all the constraints. Furthermore, the problem is not divided into subproblems as in the ATC. Surrogate model based optimization, its solution influences by the accuracy ofthe model, and this is being investigated with another DoE. A DoE of the full factorial analysis is created and selected to study in a region near the optimal solution.In such region, the result has evidently shown to be quite accurate for almost allthe surrogate models, except for max temperature, damage and strain at the hottestregion, with the largest common impact on inner wall thickness of the space nozzle. Results of the new structure of the space nozzle have shown an improvement of mass by ≈ 50%, ≈ 15% and ≈ -4%, for the three different limitations, design spacelimit, industrial limit and industrial limit with tolerance, relative to a reference value,and ≈ 10%, ≈ 35% and ≈ 25% cheaper to manufacture accordingly to the defined producibility model.

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Livres sur le sujet "Upper stage space rocket engine"

A, Curtis L., et George C. Marshall Space Flight Center., dir. Affordable In-Space Transportation. MSFC, Ala : National Aeronautics and Space Administration, Marshall Space Flight Center, 1996.

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A 20k payload launch vehicle fast track development concept using an RD-180 engine and a Centaur upper stage. Marshall Space Flight Center, Ala : National Aeronautics and Space Administration, Marshall Space Flight Center, 1995.

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Actes de conférences sur le sujet "Upper stage space rocket engine"

« Research on the High Reliability Face Seals Used in the Upper Stage Rocket Engine LH2 and LOX Turbopumps ». Dans 55th International Astronautical Congress of the International Astronautical Federation, the International Academy of Astronautics, and the International Institute of Space Law. Reston, Virigina : American Institute of Aeronautics and Astronautics, 2004. http://dx.doi.org/10.2514/6.iac-04-s.3.04.

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Gradl, Paul R., et Peter Valentine. « Carbon-Carbon Nozzle Extension Development in Support of In-space and Upper-Stage Liquid Rocket Engines ». Dans 53rd AIAA/SAE/ASEE Joint Propulsion Conference. Reston, Virginia : American Institute of Aeronautics and Astronautics, 2017. http://dx.doi.org/10.2514/6.2017-5064.

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Lee, Sun-Kyung, Sang Yeop Han, Dong-Soon Shin et R. Ray Taghavi. « Thermal-Fluidic Numerical Analysis for the Development of Heat Exchanger in the Propellant Tank Pressurization System of KSLV-II Upper Stage ». Dans ASME/JSME/KSME 2015 Joint Fluids Engineering Conference. American Society of Mechanical Engineers, 2015. http://dx.doi.org/10.1115/ajkfluids2015-03742.

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KSLV-II (Korea Space Launch Vehicle - II) launch vehicle is a three staged satellite launch vehicle using a liquid propellant propulsion system in all three stages. It will deliver 1,500 kg satellite to Sun Synchronous Orbit (SSO, 700 km, 98.2°) or 2,600 kg satellite to Low Earth Orbit (LEO, 300 km, 80.3°). Propellants for KSLV-II are kerosene as a fuel and liquid oxygen as an oxidizer for propelling. Those fuel and oxidizer are stored in on-board tanks separately. To run a liquid propellant rocket engine on ground or in flight, those propellants should be supplied to LRE’s using so-called Propellant Pressurizing Sub-system, which makes propellants be pressurized in tanks using pressurant. A pressurant for PPSS of KSLV-II is helium, which is stored in tanks located in an oxidizer tank. The stored He is under cryogenic condition (50 K) as gaseous state. Such He is heated and expanded through heat exchanger, which is using a combustion gas coming out from gas generator for turbo-pump as an energy source, to be used as pressurant. This paper contains the results of performance analysis and thermal-fluidic numerical analysis to develop the above-mentioned heat exchanger for KSLV-II upper stage (the 2nd stage). The technical requirements for such heat exchanger are as follows: pressurant mass flow rate for oxidizer tank - 0.127 kg/sec; and for fuel tank - 0.043 kg/sec. The outlet temperature of He from heat exchanger is 550±10 K.

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Davidson, D. P., et A. K. Finke. « The Design and Fabrication of a Small Highly Instrumented Counterrotating Turbine Rig ». Dans ASME 1993 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1993. http://dx.doi.org/10.1115/93-gt-396.

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The paper describes some of the design solutions that were adopted in the realization of a small two shaft, counterrotating turbine rig. The turbine, for a small space engine technology upper stage liquid propellant rocket turbo pump system, was only 100mm tip diameter with a correspondingly small annulus height. Detailed flow measurements were provided by 3 axis actuators at inlet and outlet of the working section. Torque and Power were measured on each shaft using Radio Frequency (R.F.) telemetry torquemeters. A modular design allowed the rig to be rebuilt rapidly in one and two stage configurations with different blade sets. This enabled the effect of changing turbine configurations to be studied.

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Hall, Joshua, Carl Hartsfield, Joseph Simmons et Richard Branam. « Optimized Dual-Expander Aerospike Nozzle Upper Stage Rocket Engine ». Dans 49th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition. Reston, Virigina : American Institute of Aeronautics and Astronautics, 2011. http://dx.doi.org/10.2514/6.2011-419.

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Sack, William, Julie Watanabe, Masahiro Atsumi et Hidemasa Nakanishi. « Development Progress of MB-XX Cryogenic Upper Stage Rocket Engine ». Dans 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit. Reston, Virigina : American Institute of Aeronautics and Astronautics, 2003. http://dx.doi.org/10.2514/6.2003-4486.

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Hohn, Oliver M., et Ali Guelhan. « Impact of Retro Rocket Plumes on Upper Stage Aerodynamics During Stage Separation ». Dans 20th AIAA International Space Planes and Hypersonic Systems and Technologies Conference. Reston, Virginia : American Institute of Aeronautics and Astronautics, 2015. http://dx.doi.org/10.2514/6.2015-3679.

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Durteste, Stéphane. « A Transient Model of the VINCI Cryogenic Upper Stage Rocket Engine ». Dans 43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit. Reston, Virigina : American Institute of Aeronautics and Astronautics, 2007. http://dx.doi.org/10.2514/6.2007-5531.

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« LOX/CH4 Expander Upper Stage Engine ». Dans 55th International Astronautical Congress of the International Astronautical Federation, the International Academy of Astronautics, and the International Institute of Space Law. Reston, Virigina : American Institute of Aeronautics and Astronautics, 2004. http://dx.doi.org/10.2514/6.iac-04-s.1.03.

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Kudija, Charles T., et Patrick E. Frye. « Integrated Solar Upper Stage (ISUS) engine ground demonstration (EGD) ». Dans Space technology and applications international forum - 1998. AIP, 1998. http://dx.doi.org/10.1063/1.54818.

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