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1
Wang, Yu Fei, Gong Chen, and Li Lin Han. "THE Comprehensive Survey for the Numerical Simulation of the 4th Generation Rocket Ejection Seat Thrust Vector Control System." Applied Mechanics and Materials 551 (May 2014): 523–29. http://dx.doi.org/10.4028/www.scientific.net/amm.551.523.
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The objective of this work was to develop and adapt existing computational methodologies for 3D comprehensive analysis of ejection seat aerodynamics, including rocket plume effects.Various methods were investigated for prescribing boundary and initial conditions for the seat rockets. The selected method utilizes a model that prescribes 3D nozzle exit boundary profiles extracted from detailed rocket nozzle calculations. Also a multi-domain gridding method that allows for many-to-one interface meshing was developed and tested for efficient and accurate rocket plume resolution within the 3D ejection seat computational environment. Validations were performed for Aerojet’s Pintle Escape Propulsion system (PEPS) rockets. Three dimensional ejection seat calculations with rocket power were also made to demonstrate the feasibility of the approach and the potential use of the model.
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Haw, Stephen G. "Cathayan Arrows and Meteors: The Origins of Chinese Rocketry." Journal of Chinese Military History 2, no. 1 (2013): 28–42. http://dx.doi.org/10.1163/22127453-12341243.
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Abstract Although it is now generally accepted that the rocket was invented in China, there is little agreement about exactly when this occurred. Various conflicting claims have been made, usually on the basis of dubious evidence. This article examines these claims and rejects all of them. It also discounts the hypothesis that the rocket developed from a kind of firework called a “fire-rat.” Instead, it suggests that the rocket very probably developed from Chinese fire-arrows, which carried charges of incendiary gunpowder. These became rocket-assisted arrows, fired from bows or, very often, arbalests. They achieved great ranges, of the order of two miles. They were used by the Mongols during their conquests in the mid-thirteenth century and may have formed part of the weaponry of the Chinese Song dynasty at an earlier period. The earliest reference to what were very probably true rockets, launched by their own propulsion, dates from 1272. It occurs in an account of events that took place during the siege of Xiangyang by Mongol forces. These rockets were used for signaling. There is no evidence that self-launching rockets were utilized as weapons in the mid-thirteenth century. Rockets as weapons are unlikely to have been developed any earlier than the late 1200s.
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Bolivar, Nelson Enrique, and Ivaylo T. Vasilev. "Non-Combustion ⁴He Powered Propulsion." European Journal of Engineering and Technology Research 6, no. 2 (February 16, 2021): 101–6. http://dx.doi.org/10.24018/ejers.2021.6.2.2283.
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One of the biggest hurdles nowadays rocket propulsion is the large use of fuel. The amount of fuel and the burning efficiency defines how long the rocket engine can work which intimately limits the range and the load capacity of the rockets and spaceships. This according to the Newton third law is unavoidable - in order to move forward you need to leave something behind. There have been several attempts in the past to create an engine which doesn't use fuel in the common sense, like the M drive, but so far all of them were unsuccessful. In this article we attempt to explore a novel principle, a recycling cycle of fuel, by optimizing parametrically a system that uses ⁴He phase transition.
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Freiherr, Greg. "The Little Rocket Engine That Could." Mechanical Engineering 138, no. 08 (August 1, 2016): 32–37. http://dx.doi.org/10.1115/1.2016-aug-1.
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This paper highlights advancements of a research team in the field of miniature spacecrafts and development and advantages of CubeStats. CubeSats are space age hitchhikers, that is, miniature spacecraft that fly into orbit aboard rockets whose primary payloads are full-size satellites. Paulo Lozano and his team at MIT’s Space Propulsion Lab have developed a unique kind of rocket engine for these microsatellites. The trick to building a successful ion electrospray propulsion system is to increase thrust density by jamming together as many emitters as possible. Electrospray engines also differ greatly from another form of ion propulsion, plasma ion, which also eschews chemical combustion for the efficiency of the electron. A big advantage, when asking for permission to hitchhike a ride into orbit, is that ion electrospray propulsion engines cannot explode and destroy a rocket’s primary payload. A second advantage of ion engines is their modularity and, consequently, scalability.
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Apel, Uwe, Alexander Baumann, Christian Dierken, and Thilo Kunath. "AQUASONIC – A Sounding Rocket Based on Hybrid Propulsion." Applied Mechanics and Materials 831 (April 2016): 3–13. http://dx.doi.org/10.4028/www.scientific.net/amm.831.3.
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The AQUASONIC project is aimed to develop a sounding rocket including a hybrid propulsion system based on the propellant combination nitrous oxide and polyethylene. It takes place in the frame of the STERN (Student Experimental Rockets) programme founded by the German Space Agency (DLR) in order to promote students in the area of launch vehicles. Main element of the project is the AQUASONIC rocket, which shall reach a flight altitude of 5-6 km and a velocity of MACH 1. All major activities like design, manufacturing, verification and, finally, the launch campaign will be performed by students. The rocket shall be launched at Esrange Space Centre (Sweden) in 2016. Thus, students are able to apply their skills and knowledge to a real project like it is conducted by the space industry or research organisations.
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Santos, L. M. C., L. A. R. Almeida, A. M. Fraga, and C. A. G. Veras. "EXPERIMENTAL INVESTIGATION OF A PARAFFIN BASED HYBRID ROCKET." Revista de Engenharia Térmica 5, no. 1 (July 31, 2006): 08. http://dx.doi.org/10.5380/reterm.v5i1.61658.
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Hybrid rockets are known to be simpler, safer, environmentally friend, and, more importantly, cheaper than most of the technologies for propulsion devices used today. Hybrid rockets can be applied as the propulsion system in satellites launch vehicles, micro-satellites and tactical missiles. This paper deals with combustion of ultra-high molecular weight polyethylene (UHMWPE) and paraffin as the solid fuels burning with gaseous oxygen (GOX) as well as N O as the oxidizer in lab scale hybrid rocket motors. A test 2 stand was built to carry out the experiments. The main objectives were to investigate the ignition of the solid fuels, burning performance and regression rates for different operating conditions. With paraffin-based fuel the hybrid motor had the regression rate enhanced two to three folds compared to the UHMWPE, as reported in the literature. The overall performance of the motor, with paraffin as the fuel, is comparable to other technologies. Paraffin-based hybrid rockets can, then, be a safer and cheaper alternative to satellite launch vehicles for the Brazilian space program.
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Okninski, Adam, Pawel Surmacz, Bartosz Bartkowiak, Tobiasz Mayer, Kamil Sobczak, Michal Pakosz, Damian Kaniewski, Jan Matyszewski, Grzegorz Rarata, and Piotr Wolanski. "Development of Green Storable Hybrid Rocket Propulsion Technology Using 98% Hydrogen Peroxide as Oxidizer." Aerospace 8, no. 9 (August 24, 2021): 234. http://dx.doi.org/10.3390/aerospace8090234.
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This paper presents the development of indigenous hybrid rocket technology, using 98% hydrogen peroxide as an oxidizer. Consecutive steps are presented, which started with interest in hydrogen peroxide and the development of technology to obtain High Test Peroxide, finally allowing concentrations of up to 99.99% to be obtained in-house. Hydrogen peroxide of 98% concentration (mass-wise) was selected as the workhorse for further space propulsion and space transportation developments. Over the course nearly 10 years of the technology’s evolution, the Lukasiewicz Research Network—Institute of Aviation completed hundreds of subscale hybrid rocket motor and component tests. In 2017, the Institute presented the first vehicle in the world to have demonstrated in-flight utilization for 98% hydrogen peroxide. This was achieved by the ILR-33 AMBER suborbital rocket, which utilizes a hybrid rocket propulsion as the main stage. Since then, three successful consecutive flights of the vehicle have been performed, and flights to the Von Karman Line are planned. The hybrid rocket technology developments are described. Advances in hybrid fuel technology are shown, including the testing of fuel grains. Theoretical studies and sizing of hybrid propulsion systems for spacecraft, sounding rockets and small launch vehicles have been performed, and planned further developments are discussed.
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Heeg, Francesca, Lukas Kilzer, Robin Seitz, and Enrico Stoll. "Design and Test of a Student Hybrid Rocket Engine with an External Carbon Fiber Composite Structure." Aerospace 7, no. 5 (May 13, 2020): 57. http://dx.doi.org/10.3390/aerospace7050057.
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The development of hybrid rockets offers excellent opportunities for the practical education of students at universities due to the high safety and relatively low complexity of the rocket propulsion system. During the German educational program Studentische Experimental-Raketen (STERN), students of the Technische Universität Braunschweig obtain the possibility to design and launch a sounding rocket with a hybrid engine. The design of the engine HYDRA 4X (HYbridDemonstrations-RaketenAntrieb) is presented, and the results of the first engine tests are discussed. The results for measured regression rates are compared to the results from the literature. Furthermore, the impact of the lightweight casing material carbon fiber-reinforced plastic (CFRP) on the hybrid engine mass and flight apogee altitude is examined for rockets with different total impulse classes (10 to 50 kNs). It is shown that the benefit of a lightweight casing material on engine mass decreases with an increasing total impulse. However, a higher gain on apogee altitude, especially for bigger rockets with a comparable high total impulse, is shown.
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Zaraini, Fairul Azmin, Tengku Farah Wahida Ku Chik, Nor Hafizah Abdullah, and Ahmad Ammar. "Suggestions for a Roadmap towards Becoming a Launch Capable Nation." Applied Mechanics and Materials 225 (November 2012): 561–65. http://dx.doi.org/10.4028/www.scientific.net/amm.225.561.
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Propulsion Technology Program under the National Space Agency (ANGKASA) was commenced in 2009 with an ambition to launch satellites into Low Earth Orbit (LEO) using its own independent launch vehicle. Four members of the Space Application and Technology Development (SATD) with various backgrounds have been entrusted to draft roadmap for National Satellite Launcher and at the same time conducting Research and Development (R&D) related to rocketry. The first program was solid rocket development between ANGKASA and Universiti Teknologi Malaysia (UTM) through budget allocated in Rancangan Malaysia ke-9 (RMK-9). The rockets developed in this project have been successfully launched at eastern coast of peninsula Malaysia in 2010. This achievement needs proper and effective continuation towards enabling Malaysia to be a launch capable nation. Therefore, this paper investigates rocket development programs and activities ran by various countries which could be adopted into national programs in order to spur participation in rocket science and space industries, hence materialise completion of Malaysian own launch vehicle in timely manner. Moreover, this paper will look over obstacles and potencies of rocket development with current Malaysian environment.
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Asraff, A. K., S. Sheela, Krishnajith Jayamani, S. Sarath Chandran Nair, and R. Muthukumar. "Material Characterisation and Constitutive Modelling of a Copper Alloy and Stainless Steel at Cryogenic and Elevated Temperatures." Materials Science Forum 830-831 (September 2015): 242–45. http://dx.doi.org/10.4028/www.scientific.net/msf.830-831.242.
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High performance rockets are developed using cryogenic technology. High thrust cryogenic rocket engines operating at elevated temperatures and pressures are the backbone of such rockets. The thrust chamber of such engines, which produce the thrust for the propulsion of the rocket, can be considered as structural elements. Often double walled construction is employed for these chambers for better cooling and enhanced performance. The thrust chamber investigated here has its hot inner wall fabricated out of a high conductivity high ductility copper alloy and outer wall made of a ductile stainless steel. The engine is indigenously designed and developed by ISRO and is undergoing hot tests. Inner wall is subjected to high thermal and pressure loads during operation of engine due to which it will be in the plastic regime. Evaluation of tensile properties of the copper alloy and stainless steel up to fracture, at cryogenic, ambient and elevated temperatures in parent metal and welded forms is of paramount importance for its constitutive modelling and thermo structural analysis of the thrust chamber.
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Zhang, Fan, Huiqiang Zhang, and Bing Wang. "Conceptual study of a dual-rocket-based-combined-cycle powered two-stage-to-orbit launch vehicle." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 232, no. 5 (May 1, 2017): 944–57. http://dx.doi.org/10.1177/0954410017703148.
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The liquid oxygen/methane staged cycle liquid-rocket engine is one of the most potential rocket engines in the future for its higher performance, higher fuel density and reusable capacity. Two working states of this liquid-rocket engine named as full-load state and half-load state are defined in this paper. Based on this liquid-rocket engine, a dual-rocket-based-combined-cycle propulsion system with liquid oxygen /air/methane as propellants is therefore proposed. The dual-rocket-based-combined-cycle system has then five working modes: the hybrid mode, pure ejector mode, ramjet mode, scramjet mode and pure rocket mode. In hybrid mode, the booster and ejector rockets driven by the full-load liquid-rocket engine work together with the purpose of reducing thrust demand on ejector rocket. In scramjet mode, the fuel-rich burned hot gas generated by the half-load liquid-rocket engine is used as fuel, which is helpful to reduce the technical difficulty of scramjet in hypersonic speed. The five working modes of dual-rocket-based-combined-cycle are highly integrated based on the full- or half-load state of the liquid oxygen/methane staged cycle liquid-rocket engine, and the unified single type fuel of liquid methane is adopted for the whole modes. Then a preliminary design of a horizontal takeoff two-stage-to-orbit launch vehicle is conducted based on the dual-rocket-based-combined-cycle propulsion system. Under an averaged baseline thrust and specific impulse, the launch trajectory to reach a low Earth orbit at 100 km is optimized via the pseudo-spectral method subject to maximizing the payload mass. It is shown that the two-stage-to-orbit vehicle based on the dual-rocket-based-combined-cycle can achieve the payload mass fraction of 0.0469 and 0.0576 for polar mission and equatorial mission, respectively. Conclusively, insights gained in this paper can be usefully applied to a more detailed design of the dual-rocket-based-combined-cycle powered two-stage-to-orbit launch vehicle.
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Migliorino, Mario Tindaro, Daniele Bianchi, and Francesco Nasuti. "Numerical Simulations of the Internal Ballistics of Paraffin–Oxygen Hybrid Rockets at Different Scales." Aerospace 8, no. 8 (August 5, 2021): 213. http://dx.doi.org/10.3390/aerospace8080213.
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Hybrid rockets are considered a promising future propulsion alternative for specific applications to solid or liquid rockets. In order to raise their technology readiness level, it is important to perform predictive numerical simulations of their internal ballistics. The objective of this work is to describe and validate a numerical approach based on Reynolds-averaged Navier–Stokes simulations with sub-models for fluid–surface interaction, radiation, chemistry, and turbulence. Particular attention is given to scale effects by considering two different paraffin–oxygen hybrid rocket engines and a simplified grain evolution approach from the initial to the final port diameter. Moreover, a mild sensitivity of the computed regression rate to paraffin’s melting temperature, surface radiation emissivity, and Schmidt numbers is observed. Results highlight the increasing importance of radiation effects at larger scales and pressures. A numerical rebuilding of regression rate and pressure is obtained with simulations at the time-space-averaged port diameter, producing a reasonable agreement with the available experimental data, but a noticeable improvement is obtained by considering the grain evolution in time.
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Miller, T., and A. M. Birk. "A Re-Examination of Propane Tank Tub Rockets Including Field Trial Results." Journal of Pressure Vessel Technology 119, no. 3 (August 1, 1997): 356–64. http://dx.doi.org/10.1115/1.2842316.
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When a tank containing a pressure-liquefied gas fails, one mode of failure involves the tank being propelled large distances by the released two-phase material. This mode of failure is called a tub rocket and it can pose a severe hazard to the public because of its unpredictability. Field tests were recently conducted to study the effect of explosive devices on propane tanks. The tests included tanks of various sizes up to 2000 L (500 gal).In most cases, the tests resulted in punctured tanks with transient two-phase jet releases. In some cases, the jet releases were sufficient to propel the tanks over considerable distances. In a small number of tests involving 470-L tanks, the explosive device resulted in the clean removal of a tank end, and this resulted in near-ideal launches of tub rockets. In one case, the rocket was launched vertically, and in another, the rocket was launched near 45 deg elevation angle giving a tub range of 370 m. In other cases, the explosive devices resulted in punctures, and in some of these, the resulting two-phase jet propelled the tanks over considerable distance. These examples gave a good opportunity to re-examine tub rocket models for tanks containing liquefied gases. This paper describes a theoretical model involving two-phase critical flow propulsion of cylindrical tanks. Three different critical flow models are compared, including the homogeneous equilibrium model (HEM), the homogeneous frozen model (HFM), and the Henry-Fauske model (HFK). Range predictions are compared with existing data and a model previously developed. Model predictions are calibrated to the field trial results described in the foregoing and then used to predict realistic ranges for various sizes of storage and transport tanks.
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Kim, Youngin, and Jeongho Cho. "Surface Erosion Analysis for Thermal Insulation Materials of Graphite and Carbon–Carbon Composite." Applied Sciences 9, no. 16 (August 13, 2019): 3323. http://dx.doi.org/10.3390/app9163323.
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A rocket uses fuel and oxidizers to generate propulsion by combustion and ejection, and is used for space exploration aircrafts, weapons, and satellite launches. In particular, the nozzle generating thrust of solid-propellant rockets is exposed to a high-temperature and high-pressure environment with erosion occurring from the combustion gas. When erosion occurs on the nozzle throat of such a rocket, it has a great impact on the flight performance such as reaching distance and flight speed. Many studies have been conducted to characterize erosion based on the thermochemical erosion model, since it has become important to choose nozzle materials suitable for such environments having robustness against combustion gasses of high temperature and high pressure. However, there is a limit to fully analyze the erosion characteristics only by the thermochemical erosion model. In this paper, we thus consider the mechanical erosion model with the thermochemical model for better understanding of erosion characteristics and investigate the thermochemical and mechanical erosion characteristics of nozzle throat heat-resistant materials made of graphite and carbon–carbon composites; the main factors affecting erosion are discussed by comparing the results of the experimental and theoretical models.
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Carmicino, Carmine. "Special Issue “Advances in Hybrid Rocket Technology and Related Analysis Methodologies”." Aerospace 6, no. 12 (November 26, 2019): 128. http://dx.doi.org/10.3390/aerospace6120128.
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Hybrid rockets are chemical propulsion systems that, in the most common configuration, employ a liquid oxidizer (or gaseous in much rarer cases) and a solid fuel; the oxidizer, stored in tanks, is properly injected in the combustion chamber where the solid fuel grain is bonded [...]
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Costa, Fernando S., and Gustavo A. A. Fischer. "Propulsion and Thermodynamic Parameters of van der Waals Gases in Rocket Nozzles." International Journal of Aerospace Engineering 2019 (August 14, 2019): 1–11. http://dx.doi.org/10.1155/2019/3139204.
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Propellants or combustion products can reach high pressures and temperatures in advanced or conventional propulsion systems. Variations in flow properties and the effects of real gases along a nozzle can become significant and influence the calculation of propulsion and thermodynamic parameters used in performance analysis and design of rockets. This work derives new analytical solutions for propulsion parameters, considering gases obeying the van der Waals equation of state with specific heats varying with pressure and temperature. Steady isentropic one-dimensional flows through a nozzle are assumed for the determination of specific impulse, characteristic velocity, thrust coefficient, critical flow constant, and exit and throat flow properties of He, H2, N2, H2O, and CO2 gases. Errors of ideal gas solutions for calorically perfect and thermally perfect gases are determined with respect to van der Waals gases, for chamber temperatures varying from 1000 to 4000 K and chamber pressures from 5 to 35 MPa. The effects of covolumes and intermolecular attraction forces on flow and propulsion parameters are analyzed.
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Ponomarov, Alexander Nikolaevich. "GROUND-BASED EXPERIMENTAL TESTING OF ELEMENTS OF AUTOMATION OF PNEUMATIC-HYDRAULIC SYSTEMS OF ROCKET AND SPACE TECHNOLOGY." Journal of Rocket-Space Technology 27, no. 4 (December 30, 2019): 58–61. http://dx.doi.org/10.15421/451909.
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The design and production of sophisticated technical systems, which include modern rockets and other aircraft, requires their reliability and trouble-free operation. To achieve the required level of reliability of aerospace products, a wide variety of test methods are applied at all stages of the life cycle. One of the most important systems of the launch vehicle is the pneumatic hydraulic power system of the liquid rocket propulsion system. Development of new and improvement of existing methods of control and diagnostics is one way of increasing the design and technological reliability of products of aviation and space technology. The use of functional diagnostics systems for bench and flight tests significantly increases the reliability and efficiency of space rocket technology. Researches are directed on increase of a level of reliability of products of aerospace branch. Application of systems of functional diagnostics is described at bench tests. The results of experimental researches of elements of automatics of pneumatic hydraulic power supply systems of liquid rocket engines are considered. The technique of processing of experimental data of a pulsing-acoustic method of diagnostics with use of the mathematical technology of recognition of images is presented. Deciding rules of recognition of a technical condition of object of diagnosing by results of tests are resulted. The developed method with a high degree of accuracy allows to determine the technical condition of the object of diagnosis as defective or to detect the presence of characteristic defects. Experimental testing and the proposed method of processing the results showed the efficiency of the method.
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Rao, B. N., D. Jeyakumar, K. K. Biswas, S. Swaminathan, and E. Janardhana. "Rigid body separation dynamics for space launch vehicles." Aeronautical Journal 110, no. 1107 (May 2006): 289–302. http://dx.doi.org/10.1017/s0001924000013166.
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Abstract This paper presents a systematic formulation for the simulation of rigid body dynamics, including the short period dynamics, inherent to stage separation and jettisoning parts of a satellite launcher. This also gives a review of various types of separations involved in a launch vehicle. The problem is sufficiently large and complex; the methodology involves iterations at successively lower levels of abstraction. The best choice to tackle such problems is to use state-of-the-art programming technique known as object oriented programming. The necessary classes have been identified to represent various entities in the launch vehicle separation process (e.g., gravity, aerodynamics, propulsion and separation mechanisms etc.). Simple linkages are modelled with suitable objects. This approach helps the designer to simulate a launch vehicle separation dynamics and also to analyse separation system performance. To examine the influence of the design variables on the separating bodies, statistical analyses have been performed on the upper stage separation process and pull out of ongoing stage nozzle from the spent stage of a multistage rocket carrier using retro rockets.
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Фролов, Виктор Петрович, Галина Ивановна Сокол, and Владислав Юрьевич Котлов. "ВОЛНОВОЙ ПАРАМЕТР КАК КРИТЕРИЙ В ОСНОВЕ МЕТОДА ИССЛЕДОВАНИЯ АКУСТИЧЕСКИХ ИСТОЧНИКОВ ПРИ СТАРТЕ РАКЕТ." Aerospace technic and technology, no. 3 (June 27, 2018): 4–12. http://dx.doi.org/10.32620/aktt.2018.3.01.
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The purpose of this work is to develop a method for determining the types of acoustic sources of radiation and their acoustic fields during the space rockets launch in the first seconds based on the wave parameter values. The main noise source during the space rocket launch is its propulsion system (PS). The cross-section of the nozzle is taken as the oscillation source. The theory of siren sound emission is based on the acoustic power calculation of a jet as a volume sound radiator or a radiator with a space velocity. In the model of a volumetric spherical radiator, the front of a spherical wave is a spherical surface, and the sound rays, according to the definition of the wave front, coincide with the radii of the sphere. As a result of the divergence of waves, the sound intensity decreases with distance from the source. The present work has a prospective character for clarifying the nature of the acoustic fields and for calculating the noise levels from the space rocket launch when designing the cosmodromes. In the requirements for the construction of such structures, the noise impact on the environment of infrasound radiation upon launching launch vehicles is identified. A method for determining the types of acoustic radiation sources during the space rocket launch and their acoustic fields has been developed. The method makes it possible to develop physical models of acoustic fields and apply known mathematical models to calculating their characteristics. The method is applicable for the study of acoustic emissions in the first seconds of the space rocket launch based on the determination of the wave parameter kR and allows us to provide valid data on the levels of sound pressure, intensity and acoustic power at specific points of airspace around the PS in the first seconds of the launch. The character of the acoustic wave radiation from a hole in a specific size gas flue has been studied. To calculate the acoustic characteristics, an algorithm and a program on Java programming language have been developed. Two models of acoustic field generation in the environment are described during the work of a rocket as a plane radiator and spherical waves, depending on the value of the wave parameter kR. A technique for calculating the noise of a remote control in the range for the first 1.5-4 seconds of the space rocket start time is developed
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Kindracki, Jan, Krzysztof Wacko, Przemysław Woźniak, Stanisław Siatkowski, and Łukasz Mężyk. "Influence of Gaseous Hydrogen Addition on Initiation of Rotating Detonation in Liquid Fuel–Air Mixtures." Energies 13, no. 19 (September 30, 2020): 5101. http://dx.doi.org/10.3390/en13195101.
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Hydrogen is the most common molecule in the universe. It is an excellent fuel for thermal engines: piston, turbojet, rocket, and, going forward, in thermonuclear power plants. Hydrogen is currently used across a range of industrial applications including propulsion systems, e.g., cars and rockets. One obstacle to expanding hydrogen use, especially in the transportation sector, is its low density. This paper explores hydrogen as an addition to liquid fuel in the detonation chamber to generate thermal energy for potential use in transportation and generation of electrical energy. Experiments with liquid kerosene, hexane, and ethanol with the addition of gaseous hydrogen were conducted in a modern rotating detonation chamber. Detonation combustion delivers greater thermal efficiency and reduced NOx emission. Since detonation propagates about three orders of magnitude faster than deflagration, the injection, evaporation, and mixing with air must be almost instantaneous. Hydrogen addition helps initiate the detonation process and sustain continuous work of the chamber. The presented work proves that the addition of gaseous hydrogen to a liquid fuel–air mixture is well suited to the rotating detonation process, making combustion more effective and environmentally friendly.
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Sheptun, Yuri Dmitrievich, and Sergey Viktorovich Spirkin. "CONTROL WITH REDUCING OF DISTURBING FACTORS." Journal of Rocket-Space Technology 27, no. 4 (December 30, 2019): 109–18. http://dx.doi.org/10.15421/451916.
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The structural and dynamic features of the space (moving outside the dense layers of the atmosphere) stages of rockets - carriers of spacecraft as control objects are analyzed. The reasons are investigated - disturbing factors that generate external forces and moments that determine the disturbed motion of space rocket stages. For space rocket stages, disturbing factors are: mass asymmetry of the stage relative to its longitudinal axis and angle of mismatch of the line of action of the thrust vector of the propulsion system of the stage with the longitudinal axis of the stage. It is shown that when using the stage control deviating in the hinge of the marching engine as the executive organs of the control system, the effect of auto-reduction of the mentioned disturbing factors arises. The consequence of the autocompensation of disturbing factors is the reduction of disturbing forces and moments that violate the programmed motion of the step in the pitch and yaw planes. Mass asymmetry and the angle of mismatch of the line of action of the thrust vector of its engine and the longitudinal axis of magnitude are constant. Therefore, a decrease in perturbing forces and moments is accompanied by a decrease in the amount of energy (fuel) spent on processing (zeroing) perturbations of the parameters of the perturbed motion of the stage. It is shown that if the thrust of a space-stage engine is 8000 kgf, the engine operating time (flight time of the stage) is 500 sec, the specific engine thrust is 330 sec, the mass asymmetry is 0.05 m, the angle of mismatch is 0.25 degrees, then fuel economy can reach 200 kgf. The studies were performed using mathematical modeling methods.
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Khatry, Jivan, and Fatih Aydogan. "Modeling Loss-of-Flow Accidents and Their Impact on Radiation Heat Transfer." Science and Technology of Nuclear Installations 2017 (2017): 1–15. http://dx.doi.org/10.1155/2017/1345938.
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Long-term high payload missions necessitate the need for nuclear space propulsion. The National Aeronautics and Space Administration (NASA) investigated several reactor designs from 1959 to 1973 in order to develop the Nuclear Engine for Rocket Vehicle Application (NERVA). Study of planned/unplanned transients on nuclear thermal rockets is important due to the need for long-term missions. In this work, a system model based on RELAP5 is developed to simulate loss-of-flow accidents on the Pewee I test reactor. This paper investigates the radiation heat transfer between the fuel elements and the structures around it. In addition, the impact on the core fuel element temperature and average core pressure was also investigated. The following expected results were achieved: (i) greater than normal fuel element temperatures, (ii) fuel element temperatures exceeding the uranium carbide melting point, and (iii) average core pressure less than normal. Results show that the radiation heat transfer rate between fuel elements and cold surfaces increases with decreasing flow rate through the reactor system. However, radiation heat transfer decreases when there is a complete LOFA. When there is a complete LOFA, the peripheral coolant channels of the fuel elements handle most of the radiation heat transfer. A safety system needs to be designed to counteract the decay heat resulting from a post-LOFA reactor scram.
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Jellinghaus, Katharina, Charlotte Scherer, Edouard Stauffer, Petra Urban, Michael Bohnert, and Beat P. Kneubuehl. "Deadly injuries through recoilless anti-tank weapons while military shooting practice—two case studies from Germany and Switzerland." International Journal of Legal Medicine 134, no. 6 (April 28, 2020): 2199–204. http://dx.doi.org/10.1007/s00414-020-02301-4.
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Abstract In this casuistry, two accidents from Germany and Switzerland are presented that happened during the shot of recoilless anti-tank weapons. In both cases, the injuries led to the death of two soldiers: A 22-year-old soldier in Germany was struck by the counter mass of a so-called Davis gun which had been fired by a comrade during a firing exercise; he died from his severe injuries, especially in the abdominal part of the body. As a peculiarity of the wound morphology, it was found to be a thick-layered, metallic, gray material in the wound cavity, which corresponded to the material of the counter mass that was ejected opposite to the shooting direction. The other case took place in Switzerland, where a 24-year-old soldier was seriously injured during an exercise with portable anti-tank rockets. At the time the shot was fired, he stood behind the launcher and was hit by the propulsion jet of the rocket motor. He died as well from his severe injuries, which were located at the chest done by the gas jet and by the very high pressure. In both cases, two different causes of death were present: massive blunt violence in the first case versus a jet of hot gases of very high speed and temperature in the second case.
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Chelaru, Teodor Viorel, Valentin Pana, and Adrian Chelaru. "Modelling and Simulation of Suborbital Launcher for Testing." Applied Mechanics and Materials 555 (June 2014): 32–39. http://dx.doi.org/10.4028/www.scientific.net/amm.555.32.
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The purpose of this paper is to present some aspects regarding the computational model and technical solutions for multistage suborbital launcher for testing (SLT) used to test spatial equipment and scientific measurements. The computational model consists in numerical simulation of SLT evolution for different start conditions. The launcher model presented will be with six degrees of freedom (6DOF) and variable mass. The results analysed will be the flight parameters and ballistic performances. The discussions area will focus around the technical possibility to realize a small multi-stage launcher, by recycling military rocket motors. From technical point of view, the paper is focused on national project “Suborbital Launcher for Testing” (SLT), which is based on hybrid propulsion and control systems, obtained through an original design. Therefore, while classical suborbital sounding rockets are unguided and they use as propulsion solid fuel motor having an uncontrolled ballistic flight, SLT project is introducing a different approach, by proposing the creation of a guided suborbital launcher, which is basically a satellite launcher at a smaller scale, containing its main subsystems. This is why the project itself can be considered an intermediary step in the development of a wider range of launching systems based on hybrid propulsion technology, which may have a major impact in the future European launchers programs. SLT project, as it is shown in the title, has two major objectives: first, a short term objective, which consists in obtaining a suborbital launching system which will be able to go into service in a predictable period of time, and a long term objective that consists in the development and testing of some unconventional sub-systems which will be integrated later in the satellite launcher as a part of the European space program. This is why the technical content of the project must be carried out beyond the range of the existing suborbital vehicle programs towards the current technological necessities in the space field, especially the European one.
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Митиков, Юрий Алексеевич, Наталия Михайловна Соловьева, and Богдана Александровна Крысько. "АММИАЧНЫЕ СИСТЕМЫ НАДДУВА ТОПЛИВНЫХ БАКОВ ДВИГАТЕЛЬНЫХ УСТАНОВОК РАКЕТ-НОСИТЕЛЕЙ." Aerospace technic and technology, no. 1 (February 25, 2018): 37–42. http://dx.doi.org/10.32620/aktt.2018.1.03.
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Fuel tank pressurization systems are some of most science-intensive and expensive parts of rocket carriers. There is a lot of interest in usage of structurally simple pressurization systems, including those, which don't use working bodies from the start, such as gas-generating and chemical. The possibilities of using liquid ammonia in fuel-boost systems tanks as a working fluid had been researched. The analysis of earlier and modern systems of pressurization of tanks oxidant and hot carrier rockets whose propulsion systems use liquid oxygen and kerosene had been analysed with examination of advantages and drawbacks. The purpose of conducted research is finding effective modes and simplific design of fuel tank pressurization systems with kerosene while using as working pressure of liquid ammonia. It has long been successfully used in missile technology in stabilization systems of space vehicles. The operational characteristics of ammonia and thermal aspects of its decomposition into hydrogen and nitrogen. The rate of decomposition of ammonia satisfies requirements pressurization. The gas constant of ammonia decomposition products is 978, 2 kJ. The notable disadvantage of this working body of pressurization with respect to the pressurization systems is a large the amount of heat necessary for its decomposition. The positive role of catalysts (iron, tungsten, ruthenium). Audited thermal energy on board the launch vehicle. There are offered schemes, using the heat of the engine's torch and the heat of the solid-fuel gas generator. Recommended azide gas generators, which generate pure high-temperature nitrogen. Such a scheme increases the possibility of being pressurized with helium. The possibilities of using decomposition products ammonia to pressurize the tank with oxidizer. The high efficiency of ammonia pressurizing systems example of the first stage of the medium-sized launch vehicle
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Lin, C. H., J. T. Lin, C. H. Chen, J. Y. Liu, Y. Y. Sun, Y. Kakinami, M. Matsumura, W. H. Chen, H. Liu, and R. J. Rau. "Ionospheric shock waves triggered by rockets." Annales Geophysicae 32, no. 9 (September 16, 2014): 1145–52. http://dx.doi.org/10.5194/angeo-32-1145-2014.
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Abstract. This paper presents a two-dimensional structure of the shock wave signatures in ionospheric electron density resulting from a rocket transit using the rate of change of the total electron content (TEC) derived from ground-based GPS receivers around Japan and Taiwan for the first time. From the TEC maps constructed for the 2009 North Korea (NK) Taepodong-2 and 2013 South Korea (SK) Korea Space Launch Vehicle-II (KSLV-II) rocket launches, features of the V-shaped shock wave fronts in TEC perturbations are prominently seen. These fronts, with periods of 100–600 s, produced by the propulsive blasts of the rockets appear immediately and then propagate perpendicularly outward from the rocket trajectory with supersonic velocities between 800–1200 m s−1 for both events. Additionally, clear rocket exhaust depletions of TECs are seen along the trajectory and are deflected by the background thermospheric neutral wind. Twenty minutes after the rocket transits, delayed electron density perturbation waves propagating along the bow wave direction appear with phase velocities of 800–1200 m s−1. According to their propagation character, these delayed waves may be generated by rocket exhaust plumes at earlier rocket locations at lower altitudes.
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Heller, René, Guillem Anglada-Escudé, Michael Hippke, and Pierre Kervella. "Low-cost precursor of an interstellar mission." Astronomy & Astrophysics 641 (September 2020): A45. http://dx.doi.org/10.1051/0004-6361/202038687.
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The solar photon pressure provides a viable source of thrust for spacecraft in the solar system. Theoretically it could also enable interstellar missions, but an extremely small mass per cross section area is required to overcome the solar gravity. We identify aerographite, a synthetic carbon-based foam with a density of 0.18 kg m−3 (15 000 times more lightweight than aluminum) as a versatile material for highly efficient propulsion with sunlight. A hollow aerographite sphere with a shell thickness ϵshl = 1 mm could go interstellar upon submission to solar radiation in interplanetary space. Upon launch at 1 AU from the Sun, an aerographite shell with ϵshl = 0.5 mm arrives at the orbit of Mars in 60 d and at Pluto’s orbit in 4.3 yr. Release of an aerographite hollow sphere, whose shell is 1 μm thick, at 0.04 AU (the closest approach of the Parker Solar Probe) results in an escape speed of nearly 6900 km s−1 and 185 yr of travel to the distance of our nearest star, Proxima Centauri. The infrared signature of a meter-sized aerographite sail could be observed with JWST up to 2 AU from the Sun, beyond the orbit of Mars. An aerographite hollow sphere, whose shell is 100 μm thick, of 1 m (5 m) radius weighs 230 mg (5.7 g) and has a 2.2 g (55 g) mass margin to allow interstellar escape. The payload margin is ten times the mass of the spacecraft, whereas the payload on chemical interstellar rockets is typically a thousandth of the weight of the rocket. Using 1 g (10 g) of this margin (e.g., for miniature communication technology with Earth), it would reach the orbit of Pluto 4.7 yr (2.8 yr) after interplanetary launch at 1 AU. Simplistic communication would enable studies of the interplanetary medium and a search for the suspected Planet Nine, and would serve as a precursor mission to α Centauri. We estimate prototype developments costs of 1 million USD, a price of 1000 USD per sail, and a total of < 10 million USD including launch for a piggyback concept with an interplanetary mission.
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ARAKAWA, Yoshihiro. "Ion Propulsion Rocket." SHINKU 38, no. 6 (1995): 600–604. http://dx.doi.org/10.3131/jvsj.38.600.
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Greatrix, David, Ivett Leyva, Dario Pastrone, Valsalayam Sanal Kumar, and Michael Smart. "Chemical Rocket Propulsion." International Journal of Aerospace Engineering 2012 (2012): 1–2. http://dx.doi.org/10.1155/2012/715706.
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Al-shemmeri, T. T. "Rocket propulsion elements." Journal of Mechanical Working Technology 16, no. 2 (April 1988): 226. http://dx.doi.org/10.1016/0378-3804(88)90172-6.
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Viganò, Davide, Adriano Annovazzi, and Filippo Maggi. "Monte Carlo Uncertainty Quantification Using Quasi-1D SRM Ballistic Model." International Journal of Aerospace Engineering 2016 (2016): 1–8. http://dx.doi.org/10.1155/2016/3765796.
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Compactness, reliability, readiness, and construction simplicity of solid rocket motors make them very appealing for commercial launcher missions and embarked systems. Solid propulsion grants high thrust-to-weight ratio, high volumetric specific impulse, and a Technology Readiness Level of 9. However, solid rocket systems are missing any throttling capability at run-time, since pressure-time evolution is defined at the design phase. This lack of mission flexibility makes their missions sensitive to deviations of performance from nominal behavior. For this reason, the reliability of predictions and reproducibility of performances represent a primary goal in this field. This paper presents an analysis of SRM performance uncertainties throughout the implementation of a quasi-1D numerical model of motor internal ballistics based on Shapiro’s equations. The code is coupled with a Monte Carlo algorithm to evaluate statistics and propagation of some peculiar uncertainties from design data to rocker performance parameters. The model has been set for the reproduction of a small-scale rocket motor, discussing a set of parametric investigations on uncertainty propagation across the ballistic model.
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Misterek, Dean. "Book Review: Rocket Propulsion." AIAA Journal 57, no. 7 (July 2019): 3110. http://dx.doi.org/10.2514/1.j058612.
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Gluck, Paul. "Rocket Propulsion with Sparklers." Physics Teacher 44, no. 6 (September 2006): 336–37. http://dx.doi.org/10.1119/1.2336131.
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KUBOTA, Naminosuke. "Ducted rocket propulsion technology." Journal of the Japan Society for Aeronautical and Space Sciences 39, no. 446 (1991): 99–106. http://dx.doi.org/10.2322/jjsass1969.39.99.
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Winterberg, F. "Nuclear rocket propulsion principle." Acta Astronautica 32, no. 6 (June 1994): 459–62. http://dx.doi.org/10.1016/0094-5765(94)90046-9.
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Winterberg, F. "Deuterium microbomb rocket propulsion." Acta Astronautica 66, no. 1-2 (January 2010): 40–44. http://dx.doi.org/10.1016/j.actaastro.2009.05.020.
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Rugescu, Radu D. "Conflicting Design Issues Hallmark the Compound Rocket Motors." Applied Mechanics and Materials 555 (June 2014): 49–56. http://dx.doi.org/10.4028/www.scientific.net/amm.555.49.
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A series of conflicting requirements to which the design of the compound rocket motors are subjected are considerably diminishing the efficiency of that type of rocket motors. The necessity of delivering a high thrust level at lift-off and during the sustainer phase up to optimal velocity accommodation, coupled with the requirement for a slander body configuration for air drag mitigation during the atmospheric ascent are the dominant issues that end into a compromise that lower the overall efficiency to a great extent, in comparison the available theoretical performance of compound rocket motors. The optimal design is imposing a lower mass ratio that expected and is the main cause of efficiency reduction when very high velocity requirements are searched for, like in orbital launchers. However, the margins of propulsive efficiency build-up can be conveniently manipulated through geometrical optimization and attentive risk management of the propulsion system as shown by the experimental results protruded during the development of the NERVA-ORVEAL space rocket motor.
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Basharina, T. A., M. G. Goncharov, S. N. Lymich, V. S. Levin, and D. P. Shmatov. "Low-thrust liquid-propellant rocket engines as part of advanced ultralight rocket vehicle systems." Spacecrafts & Technologies 5, no. 1 (March 25, 2021): 5–13. http://dx.doi.org/10.26732/j.st.2021.1.01.
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This work examines the most promising design solutions for the creation of propulsion systems for ultra-light launch vehicles by small private enterprises in the rocket and space industry. Comparison of the metal consumption of the combustion chambers with the energy characteristics at different operating pressures showed that the most optimal operating pressure is 12,16 MPa. Comparison of the relative and absolute values of the masses of various configurations describes the nature of the relationship between the number of combustion chambers and the total mass of the propulsion system. It was found that nine-chamber propulsion systems with cameras made with extensive use of additive technologies best meet the key requirements. The analysis carried out includes an assessment of the design parameters of both various components and assemblies and the propulsion system as a whole. Various layouts of propulsion systems are considered in detail, the required degree of technological complexity of structures of various units and assemblies, their production cost are estimated. The ratio of the obtained mass-energy characteristics was achieved through the implementation of design solutions that became available due to the use of additive technologies. The obtained results of preliminary calculations demonstrate the applicability and efficiency of design solutions considered for use in the propelled propulsion system for a promising launch vehicle.
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Colasurdo, Guido, and Lorenzo Casalino. "Energy Management in Rocket Propulsion." Journal of Propulsion and Power 16, no. 4 (July 2000): 705–8. http://dx.doi.org/10.2514/2.5630.
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Galfetti, L., L. T. De Luca, F. Severini, L. Meda, G. Marra, M. Marchetti, M. Regi, and S. Bellucci. "Nanoparticles for solid rocket propulsion." Journal of Physics: Condensed Matter 18, no. 33 (August 4, 2006): S1991—S2005. http://dx.doi.org/10.1088/0953-8984/18/33/s15.
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Halai, Chandrahas M. "How does rocket propulsion work?" Resonance 16, no. 1 (January 2011): 65–68. http://dx.doi.org/10.1007/s12045-011-0010-7.
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Martin, James A. "Selecting hydrocarbon rocket propulsion technology." Acta Astronautica 16 (January 1987): 343–55. http://dx.doi.org/10.1016/0094-5765(87)90123-8.
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Kulygin, V. М. "AN ION ROCKET PROPULSION PHENOMENOLOGY." Problems of Atomic Science and Technology, Ser. Thermonuclear Fusion 43, no. 4 (2020): 110–17. http://dx.doi.org/10.21517/0202-3822-2020-43-4-110-117.
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MOTYL, Krzysztof, Mirosław MAKOWSKI, Bogdan ZYGMUNT, Zbigniew PUZEWICZ, and Janusz NOGA. "A Concept for Striking Range Improvement of the GROM/PIORUN Man-Portable Air-Defence System." Problems of Mechatronics Armament Aviation Safety Engineering 8, no. 1 (March 31, 2017): 55–70. http://dx.doi.org/10.5604/01.3001.0009.8994.
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This paper presents a concept for striking distance performance improvement of the GROM/PIORUN Man-Portable Air-Defence System rocket missiles by increasing the rated diameter of the rocket missile propulsion system and its fuel charge weight. A mathematical and physical model of the GROM rocket missile was designed and its enhanced propulsion system was simulated in a computer environment. The computer simulation results were displayed on plot charts.
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Ivanchenko, A. M. "A paste-propellant rocket propulsion system." Kosmìčna nauka ì tehnologìâ 5, no. 4 (July 30, 1999): 3–10. http://dx.doi.org/10.15407/knit1999.04.003.
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DeLuca, Luigi Tonino. "Innovative Solid Formulations for Rocket Propulsion." Eurasian Chemico-Technological Journal 18, no. 3 (November 5, 2016): 181. http://dx.doi.org/10.18321/ectj424.
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Solid rocket propulsion enjoys several unique properties favoring its use in space exploration and military missions still for decades to come, in spite of being by far the most mature propulsion technology among those currently employed. Yet, solid rocket propellants also suffer a limited performance in terms of gravimetric specifi c impulse. Although many high-energy density materials have been identifi ed, most of them are far from being practically usable in the short range due to a variety of severe diffi culties, including cost considerations. Presently, no integrated vehicle designs make use of these potential propellant ingredients<br />and formulations. Work is continuing worldwide and a broad overview will be discussed in this paper based on a joint international editorial effort just completed. After a quick historical survey, the current situation in terms of advanced solid oxidizers, metal fuels, and binder systems is scrutinized. Particular attention is paid to Ammonium Dinitramide (ADN)-based formulations to overcome the limitations of the currently used ones based on Ammonium Perchlorate (AP). The latter imply not only a limited gravimetric specifi c impulse but also a negative impact on the environment because of copious emissions of hydrochloric acid (HCl) as well as personal health because of perchlorate competition with iodide in entering the thyroid gland. Based on recent experimental investigations, due to its intrinsic ballistic properties, it turns out that ADN-based dual-oxidizer systems with Albased dual-metal fuels and inert or energetic binders are promising solutions for a variety of solid rocket propulsion aiming respectively at minimizing environmental impact (ADN + Ammonium Nitrate AN) or maximizing performance (ADN + AP). Yet, a lot of work remains to be done in order to upgrade these formulations to industrial applications. In particular, adequate analyses of manufacture, mechanical, and hazard properties are required.
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Baldin, A. A. "Ecological aspect of launch vehicles development by criterion of minimal cost." Ecology and Noospherology 25, no. 3-4 (May 29, 2014): 114–19. http://dx.doi.org/10.15421/031427.
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One of the topical problems in modern aerospace engineering is accordance between ecological requirements and performance of the vehicle. On the other hand, problem of economical efficiency leads to change of the main criterion of designing to the minimization of costs (instead of maximal performance). According to modern trends of “low-cost” vehicles, different concepts of the future cost-effective launch vehicles are considered. It is necessary to validate these concepts according to requirements of ecological safety for the purpose of detection of the dominant launch vehicle configuration. Typical configurations of the future 'low-cost' launch vehicle are presented by 6 conceptual groups (Koelle, 2001). Conceptual group 1 (CG1) is presented by the Ballistic “Single stage to orbit” (SSTO) reusable vehicle. All vehicles which use classical rocketry scheme of the propulsion trajectory are called “Ballistic” i.e. the ballistic vehicle is lifted to orbit under the impact of rocket engines thrust. CG1-vehicle is able to reach the low earth orbit (LEO) without stage separation reducing the number of required rocket engines. Technological feasibility of SSTO concepts is proven by numerous studies (Koelle, 2001). CG2 representatives are ballistic “Two stages to orbit” (TSTO) reusable vehicles. The difference between CG1 and CG2 consists in application of vacuum rocket engines in the second stage and, consequently, stage separation. CG2 are the most mass-effective vehicles. CG3 is presented by the winged SSTO vehicles with rocket propulsion by “Lifting body” aerodynamic scheme. Ascensional force is provided by the aerodynamic shape of the vehicle’s structure at high speeds. Winged TSTO vehicles with rocket propulsion and parallel or tandem staging form the CG4. The winged configuration provides wide landing capability for both stages. CG5 is presented by winged TSTO vehicles with airbreathing propulsion in the first stage and rocket-propelled second stage. Airbreathing jet engines provide high reusability ratio comparing with other concepts as well as the widest landing capability. Aerospace Plane with scramjet-rocket propulsion forms CG6. The vehicle is able to reach near-cosmic speed in rarefied layers of the atmosphere and then accelerate with rocket engines. The most ecologically important resemblance of represented concepts is reusability. This reduces space debris formation (due to lack of waste hardware). Reusable launch vehicles can also be used to return the spent satellites. Structural differences between the concepts form 3 criterions of comparison by ecological impact: 1) propellant toxicity; 2) safety of surface facilities (vehicle damage inside the atmosphere); 3) probability of space debris formation (vehicle damage outside the atmosphere). Comparison of the concepts by these criterions allows substantiating the most ecologically acceptable direction of research. Results of the comparison demonstrate that the most ecologically acceptable low-cost launch vehicle configuration is: Ballistic SSTO or TSTO reusable launch vehicle with “LOX+LH2” propellant. The results can be explained by following way: combustion products of the propellant “liquid oxygen + liquid hydrogen” are absolutely safe for environment. It also provides maximal performance of rocket engine (due to the highest specific impulse). Ballistic ascent scheme allows using relatively simple technologies and provides high reliability level. In combination with minimal time of atmospheric flight this provides high level of safety for surface facilities. These results may be used for substantiation of dominant research direction.
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Perkins, H. Douglas, Scott R. Thomas, and James R. DeBonis. "Rocket-Based Combined Cycle Propulsion System Testing." Journal of Propulsion and Power 14, no. 6 (November 1998): 1065–67. http://dx.doi.org/10.2514/2.5375.
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Venkateswaran, Sankaran, Charles L. Merkle, and Stefan T. Thynell. "Analysis of direct solar thermal rocket propulsion." Journal of Propulsion and Power 8, no. 3 (May 1992): 541–47. http://dx.doi.org/10.2514/3.23511.
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Moon, Heejang, Seongjoo Han, Youngjun You, and Minchan Kwon. "Hybrid Rocket Underwater Propulsion: A Preliminary Assessment." Aerospace 6, no. 3 (March 6, 2019): 28. http://dx.doi.org/10.3390/aerospace6030028.
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This paper presents an attempt to use the hybrid rocket for marine applications with a 500 N class hybrid motor. A 5-port high density polyethylene (HDPE) fuel grain was used as a test-bed for the preliminary assessment of the underwater boosting device. A rupture disc preset to burst at a given pressure was attached to the nozzle exit to prevent water intrusion where a careful hot-firing sequence was unconditionally required to avoid the wet environment within the chamber. The average thrust level around 450 N was delivered by both a ground test and an underwater test using a water-proof load cell. However, it was found that instantaneous underwater thrusts were prone to vibration, which was due in part to the wake structure downstream of the nozzle exit. Distinctive ignition curves depending on the rupture disc bursting pressure and oxidizer mass flow rate were also investigated. To assess the soft-start capability of the hybrid motor, the minimum power thrust, viewed as the idle test case, was evaluated by modulating the flow controlling valve. It was found that an optimum valve angle, delivering 16.3% of the full throttle test case, sustained the minimum thrust level. This preliminary study suggests that the throttable hybrid propulsion system can be a justifiable candidate for a short-duration, high-speed marine boosting system as an alternative to the solid underwater propulsion system.