Relevant bibliographies by topics / Rocket engines – Design and construction

Academic literature on the topic 'Rocket engines – Design and construction'

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Published: 10 December 2022
Last updated: 28 January 2023

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Journal articles on the topic "Rocket engines – Design and construction"

1

Zosimovych, Nickolay. "Sounding Rocket Preliminary Design." European Journal of Engineering and Technology Research 6, no. 2 (February 23, 2021): 136–41. http://dx.doi.org/10.24018/ejers.2021.6.2.2368.

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This article aims at studying sounding rocket (SR) class, or those vehicles launched from the earth and carrying out various experiments at high altitudes. After the end of their mission, rockets should deliver a payload to a given point or to give range. In general, design tasks are divided into direct and inverse ones. This paper focuses on direct designing tasks. In this respect, the set values are maximum range, payload mass and restrictions on the sounding rocket construction. As a result, a SR general design technique is proposed. This technique includes selection of the SR initial parameters, the number of stages, the relative weights of the fuel components, the specific thrust of the engines for each stage and the initial transverse load on the SR. After selecting the fuel composition, engine design features and rocket as an object, the starting mass can be represented as a complex mathematical function. Moreover, predetermined maximum range, rocket payload, selected fuel components, structural design, materials, design parameters allow for a definite SR weight determination. Finally, overall and power characteristics of the SR are proposed within this technique. Such characteristics are accepted as design parameters.
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Zosimovych, Nickolay. "Sounding Rocket Preliminary Design." European Journal of Engineering and Technology Research 6, no. 2 (February 23, 2021): 136–41. http://dx.doi.org/10.24018/ejeng.2021.6.2.2368.

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This article aims at studying sounding rocket (SR) class, or those vehicles launched from the earth and carrying out various experiments at high altitudes. After the end of their mission, rockets should deliver a payload to a given point or to give range. In general, design tasks are divided into direct and inverse ones. This paper focuses on direct designing tasks. In this respect, the set values are maximum range, payload mass and restrictions on the sounding rocket construction. As a result, a SR general design technique is proposed. This technique includes selection of the SR initial parameters, the number of stages, the relative weights of the fuel components, the specific thrust of the engines for each stage and the initial transverse load on the SR. After selecting the fuel composition, engine design features and rocket as an object, the starting mass can be represented as a complex mathematical function. Moreover, predetermined maximum range, rocket payload, selected fuel components, structural design, materials, design parameters allow for a definite SR weight determination. Finally, overall and power characteristics of the SR are proposed within this technique. Such characteristics are accepted as design parameters.
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3

Палюх, Алексей Владимирович, Сергей Александрович Мотылев, and Леонид Прокофьевич Малый. "КОНСТРУКТОРСКОЕ И ТЕХНОЛОГИЧЕСКОЕ ОБЕСПЕЧЕНИЕ ИЗГОТОВЛЕНИЯ КОРПУСА РАКЕТНОГО ДВИГАТЕЛЯ НА ТВЕРДОМ ТОПЛИВЕ ТИПА “КОКОН”." Aerospace technic and technology, no. 8 (August 31, 2019): 22–27. http://dx.doi.org/10.32620/aktt.2019.8.04.

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Rocket engines on firm fuel are widely applied in the rocket and space-rocket technics. The case is the basic bearing element of the engine on firm fuel and carries out functions of the chamber of combustion, and also is the basic power element of the engine and simultaneously a part of a power design of the rocket. In the article on firm fuel of type "cocoon", mission and requirements the description of a design of the case of the rocket engine results in elements of a design of case rocket engines on firm fuel of type "cocoon". It is determined the applied materials and technology requirements to them. The choice of constructional materials for each separate element and knot depends on many factors and the requirements shown to a product. Properties of composite material substantially depend as on parity between binding and reinforcing components, and from parameters of the technological process of their manufacturing. It is determined a technological route of manufacturing of case rocket engines on firm fuel of type "cocoon". The technological route of manufacturing of any product appreciably depends on applied materials, manufacture type, presence of an industrial base, a design of a product, its form, technical requirements, etc. It is determined the manufacturing of an internal heat-shielding covering of case rocket engines on firm fuel of type "cocoon". It is determined the manufacturing internal heat-shielding covering of the bottoms forward, back and a cylindrical part of the case. The internal heat-shielding covering of the bottoms of case rocket engines on firm fuel of type "cocoon" represents a multilayered design with flanges, cuffs, jacks. In this connection manufacturing heat-shielding covering of the bottoms is spent separately. Before calculation heat-shielding covering of a cylindrical part on prepared mandrel covers (technological and a fixing layer), technological covers of the forward and back bottoms keep within consistently and also are established heat-shielding covering the forward and back bottoms. It is determined the manufacturing of a power cover of the case, a cover of communication and joint knots. Power cover and the cover of communication of the case make spirally-ring "wet" winding of the carbon fiber impregnated epoxy resin, on winding the machine tool. Forward and back knots of a joint of the case are made by a method of winding of layers of carbon fabric. Layers of a carbon fabric alternate with ring and spiral layers of a carbon fiber carrying out a role of consolidation of layers of carbon fabric and communication of forward and back knots of a joint.
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Золотько, Олександр Євгенович, Олена Василівна Золотько, Олександра Валеріївна Сосновська, Олександр Сергійович Аксьонов, and Ірина Сергіївна Савченко. "ОСОБЛИВОСТІ КОНСТРУКТИВНИХ СХЕМ ДВИГУНІВ З ІМПУЛЬСНИМИ ДЕТОНАЦІЙНИМИ КАМЕРАМИ." Aerospace technic and technology, no. 2 (April 27, 2020): 4–10. http://dx.doi.org/10.32620/aktt.2020.2.01.

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The pressure of the products of chemical reactions in the chamber of a rocket engine increases significantly if the rocket fuel components burn in the detonation mode. In this case, it can get to a simpler and more reliable expulsion propellant feed system instead of a turbopump feed system. The value of heat release power (MW / liter) of detonation engines is several orders of magnitude larger than that of aircraft and rocket engines operating in the Brighton cycle. The high rate of energy released in the detonation mode can significantly reduce the mass, the inertia, and overall dimensions of the propulsion system. Due to these features, detonation chambers are advisable to be used as part of ejector pulsed detonation engines, together with a turbine – in electric power generators of spacecraft, in a hybrid design – together with turbofan or turboprop engines, etc. In the article are considered various design schemes of pulse detonation engines (PDE): single-chamber and multi-chamber pulsed detonation engines; an ejector PDE system; a hybrid PDE and an integrated detonation-turbine unit with a detonation chamber in the form of a spiral and with a multi-chamber detonation device. The possibility of pulsation frequency increase is realized in the multi-chamber pulsed detonation engine, and the possibility of thrust size increase is realized in PDE with ejector. Replacing traditional chambers with detonation chambers in the construction of gas turbine jet engine will allow providing a decrease in propellant flow rate value from 8 % to 10 % on some estimations. In the hybrid detonation propulsion plant advantages inherent to the detonation cycle combine with positive features of a turbo-compressor jet engine. A combination of PDE and turbine allows creating the cogeneration propulsion system in that a turbine is used for the production of electric power, and detonation chamber – for the creation of thrust impulse. Practical realization of hybrid pulse detonation turbo-engine and the integrated detonation-turbine device is possible if two key complex problems will be solved. These problems are the detonation waves weakening on input in a turbine and the bearing and shaft necessary work resource increasing into a detonation pulsating stream
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Ricci, Daniele, Francesco Battista, and Manrico Fragiacomo. "Transcritical Behavior of Methane in the Cooling Jacket of a Liquid-Oxygen/Liquid-Methane Rocket-Engine Demonstrator." Energies 15, no. 12 (June 7, 2022): 4190. http://dx.doi.org/10.3390/en15124190.

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The successful design of a liquid rocket engine is strictly linked to the development of efficient cooling systems, able to dissipate huge thermal loads coming from the combustion in the thrust chamber. Generally, cooling architectures are based on regenerative strategies, adopting fuels as coolants; and on cooling jackets, including several narrow axial channels allocated around the thrust chambers. Moreover, since cryogenic fuels are used, as in the case of oxygen/methane-based liquid rocket engines, the refrigerant is injected in liquid phase at supercritical pressure conditions and heated by the thermal load coming from the combustion chamber, which tends to experience transcritical conditions until behaving as a supercritical vapor before exiting the cooling jacket. The comprehension of fluid behavior inside the cooling jackets of liquid-oxygen/methane rocket engines as a function of different operative conditions represents not only a current topic but a critical issue for the development of future propulsion systems. Hence, the current manuscript discusses the results concerning the cooling jacket equipping the liquid-oxygen/liquid-methane demonstrator, designed and manufactured within the scope of HYPROB-NEW Italian Project. In particular, numerical results considering the nominal operating conditions and the influence of variables, such as the inlet temperature and pressure values of refrigerant as well as mass-flow rate, are shown to discuss the fluid transcritical behavior inside the cooling channels and give indications on the numerical methodologies, supporting the design of liquid-oxygen/liquid-methane rocket-engine cooling systems. Validation has been accomplished by means of experimental results obtained through a specific test article, provided with a cooling channel, characterized by dimensions representative of HYPROB DEMO-0A regenerative combustion chamber.
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Takao, Y., T. Shibui, and S. Saito. "A Design and Construction of Civil Work for Rocket Engine Firing Test Facilities." Concrete Journal 27, no. 6 (1989): 38–46. http://dx.doi.org/10.3151/coj1975.27.6_38.

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Pîslaru-Dănescu, Lucian, Alexandru-Mihail Morega, Rareş-Andrei Chihaia, Ionel Popescu, Mihaela Morega, Lică Flore, Marius Popa, and Eros-Alexandru Pătroi. "New Type of Linear Magnetostrictive Motor Designed for Outer Space Applications, from Concept to End-Product." Actuators 10, no. 10 (October 14, 2021): 266. http://dx.doi.org/10.3390/act10100266.

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The use of the linear magnetostrictive motor (LMM) in outer space, in the absence of Earth’s gravitational field and where extreme temperatures manifest, involves innovative technical solutions that result in significant construction changes. This paper highlights these constructive changes and presents the mathematical modeling followed by the numerical simulation of different operating regimes of LMM. The novelty of the design resides in using a bias coil instead, in addition to permanent magnets, to magnetize the magnetostrictive core and pulse width modulated (PWM) power sources to control the two coils of the LMM (bias and activation). The total absorbed current is less than 2 A, which results in the reduction of Joule losses. Moreover, a PWM source is provided to power and control a set of three Peltier elements aimed at cooling the device. The experiments validate the design of the LMM, which elicits it to power and control devices that may modulate fuel injection for rocket engines or for machines used to adjust positioning on circumterrestrial orbits.
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Budur, О., S. Nikul, M. Petrushenko, V. Holovan, O. Serheyev, and S. Tarasenko. "DEVELOPMENT OF RECOMMENDATIONS FOR CALCULATING THE MECHANISM OF SEPARATION OF PERSPECTIVE SOLID PROPELLANT ROCKETS WITH A NEGATIVE IN-FLIGHT WARHEAD AND MECHANISM OF SEPARATION." Collection of scientific works of Odesa Military Academy 1, no. 12 (December 27, 2019): 81–88. http://dx.doi.org/10.37129/2313-7509.2019.12.1.81-88.

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In today's difficult international environment, especially in view of the situation in which Ukraine finds itself, it is necessary to note that each of the states with different social structures makes a titanic effort to prevent a nuclear war, to preserve and strengthen the peace. Engaging in peaceful construction, it is necessary to protect the defense capability of the Ukrainian state, where we have no right to allow the weakening for a second. The prospect of solving such problems lies to some extent on missile systems, as they are high-precision weapons. To accomplish these missiles, it is necessary to have a modern element base. Particularly important is the task of replacing already outdated samples of weapons with new, where the latest achievements of scientific and technical work, new approaches in design. Solid propellant rocket engines are now widespread. It has the following major advantages: high reliability, ease of use, constant readiness for action. Rocket launchers are used in all classes of modern military complexes. A variety of applications and tasks contribute to the development of a wide range of structures that differ in overall, mass, traction, time and other characteristics. When approaching such work, it is necessary to take into account the classification of this type of missiles, to analyze the requirements advanced to the missiles in terms of standard, operational and production and economic requirements. In addition, to select and justify the scheme of the rocket. Determine the type of start, engine rocket. Particular attention should be given to the determination of structural materials and the choice of the missile flight program. In the analysis of the requirements for the systems of separation of the warhead, it is necessary to consider their schematic diagrams, on the basis of which the choice of scheme for a prospective missile and the design of the mechanism of separation of the warhead. Based on the analysis and calculations, fast detachable devices were used as fasteners of the warhead: burst bolts having a sealed axial channel filled with pyrotechnic composition with a lighter. Of the possible three groups of separation mechanisms are selected pushing mechanisms that act on the warhead and the hull forces in the direction of the longitudinal axis of the rocket. As a pusher, the compartment used a spring pusher. Based on the calculations, we can say that the system compartment warhead can provide its secure attachment to the rocket body with the help of burst bolts. These mounting mechanisms are compact, small in weight, safe to operate and have a simple construction. As a mechanism of separation of the warhead can be selected by pushing the mechanisms of separation, namely - spring mechanisms, because they provide a reliable separation of the warhead and minimal disturbance of its movement in the separation process. As a result of the calculation of the mechanism of separation of the warhead, calculated the necessary and sufficient condition of the separation, its relative velocity after separation, found the necessary values of the force of pushing, providing reliable operation of the pushing mechanism.
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Somov, V. V. "DETERMINATION OF THE TYPE OF A SINGLE-USE GRENADE LAUNCHER BASED ON ITS COMPOSITE PARTS AND FRAGMENTS OF REACTIVE GRENADE FOUND AT THE PLACE OF ACCIDENT." Theory and Practice of Forensic Science and Criminalistics 17 (November 29, 2017): 245–52. http://dx.doi.org/10.32353/khrife.2017.31.

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In carrying out an investigation into the explosion, among others, the investigative version of the use of a single-use reactive grenade launcher is being considered. The most common for criminal explosions are applied grenade launchers RPG-18, RPG-22, RPG-26. Their use is due to a number of such properties as small size and weight, which makes it possible to transfer them covertly, the range of the shot significantly exceeding the range of the hand grenade throw, the high detonating effect of the rocket grenade explosion. The single-use rocket launchers are generally of the same design. Their differences are in the features of the components construction and dimensional characteristics, which are given in the article. On the basis of expert practice, details ofgrenade launchers that remain at the site of the explosion and have the least damage are determined. These details are the objects of investigation of the explosion technical expertise. These objects include launchers of grenade launchers and rocket parts ofjet grenades. The design features of the launchers, their dimensional characteristics and marking symbols make it possible to determine their belonging to a specific type of jet grenade launchers. Missile parts of jet grenades differ in the form of the combustion chamber of the jet engine, nozzle, in the size ofthe outlet section of the nozzle, in the form and size of the stabilizerfeathers. To determine the belonging of the rocket part of the grenade to a specific type ofjet grenade launcher, it’s necessary to establish a set of structural features and dimensional characteristics. At considerable damage of the combustion chamber of the jet engine, as a rule, the nozzle block remains intact that allows to define diameter of critical section of a nozzle, and on it to establish type of the used single-use grenade launcher.
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Wisniewski, Adam, Maciej Malicki, and Wojciech Manaj. "Visual and microscopic examination of the rocket engine combustion chamber." Aircraft Engineering and Aerospace Technology 92, no. 3 (February 18, 2019): 368–75. http://dx.doi.org/10.1108/aeat-06-2018-0166.

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Purpose This paper aims to enhance the selection of the best material of the rocket engine combustion chamber. The chamber has been destroyed during dynamometer tests, and the goal of this inspection is to verify the nature of the damage in the context of checking the usefulness of this type of graphite for the combustion chamber construction. Design/methodology/approach This paper presents the results of visual and microscopic inspection of the rocket engine combustion chamber of Ø50 × 165 mm in dimension, which was made of R type graphite. Findings An analysis of the fracture surface shows that in the inspected combustion chamber voids and inclusions are present. EDS analysis of the fracture surface shows that in the inspected combustion chamber inclusions are present which have a relatively high amount of elements like: Ti, C, S, V, Si, O and a relatively small amount of Fe and Ni. Research limitations/implications Research limitations is concerned the failure analysis by a scanning electron microscope (SEM) Zeiss EVO 25 MA with EDS detector: Brüker X Flash Detector 5010 125 eV and Espirit 1.9.0.2176 EDS software. Practical implications Designing of the engine combustion chamber the researches can select the best of the rocket engine combustion chamber, made of R type graphite, with the minimum voids and inclusions to decrease the possibility of bursting of this chamber. Originality/value The most dangerous issues in the inspected combustion chamber during an outflow are hot gases as a result of high fuel combustion temperature, so it causes the nozzle heating and the engine stress increase of visible inclusions in cross-sections.
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Dissertations / Theses on the topic "Rocket engines – Design and construction"

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Baig, Saood Saeed. "A simple moving boundary technique and its application to supersonic inlet starting /." Thesis, McGill University, 2008. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=112555.

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In this thesis, a simple moving boundary technique has been suggested, implemented and verified. The technique may be considered as a generalization of the well-known "ghost" cell approach for boundary condition implementation. According to the proposed idea, the moving body does not appear on the computational grid and is allowed to move over the grid. The impermeable wall boundary condition is enforced by assigning proper gasdynamic values at the grid nodes located inside the moving body close to its boundaries (ghost nodes). The reflection principle taking into account the velocity of the boundaries assigns values at the ghost nodes. The new method does not impose any particular restrictions on the geometry, deformation and law of motion of the moving body.
The developed technique is rather general and can be used with virtually any finite-volume or finite-difference scheme, since the modifications of the schemes themselves are not required. In the present study the proposed technique has been incorporated into a one-dimensional non-adaptive Euler code and a two-dimensional locally adaptive unstructured Euler code.
It is shown that the new approach is conservative with the order of approximation near the moving boundaries. To reduce the conservation error, it is beneficial to use the method in conjunction with local grid adaptation.
The technique is verified for a number of one and two dimensional test cases with analytical solutions. It is applied to the problem of supersonic inlet starting via variable geometry approach. At first, a classical starting technique of changing exit area by a moving wedge is numerically simulated. Then, the feasibility of some novel ideas such as a collapsing frontal body and "tractor-rocket" are explored.
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Miller, Robert E. Hartfield Roy J. "Design and testing of a gas distribution method for pulsed inductive thruster." Auburn, Ala, 2008. http://hdl.handle.net/10415/1405.

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Doan, Andrew W. "3-D flow and performance of a rocket pump inducer at design and off-design flow rates." Thesis, This resource online, 1994. http://scholar.lib.vt.edu/theses/available/etd-11242009-020251/.

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Russ, David Phillip. "Analysis of a heavy lift launch vehicle design using small liquid rocket engines." Thesis, Massachusetts Institute of Technology, 1988. http://hdl.handle.net/1721.1/35340.

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Denton, Brandon Lee. "Design and analysis of rocket nozzle contours for launching pico-satellites /." Online version of thesis, 2008. http://hdl.handle.net/1850/6003.

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Dausen, David F. "Design of a premixed gaseous rocket engine injector for ethylene and oxygen." Thesis, Monterey, Calif. : Naval Postgraduate School, 2006. http://bosun.nps.edu/uhtbin/hyperion.exe/06Dec%5FDausen.pdf.

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Thesis (M.S. in Astronautical Engineering)--Naval Postgraduate School, December 2006.
Thesis Advisor(s): Christopher M. Brophy. "December 2006." Includes bibliographical references (p. 97). Also available in print.
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Cunha, Marco Antonio Hidalgo. "A computational tool for the design and optimization of supersonic turbines with application on turbopump rocket engines." Instituto Tecnológico de Aeronáutica, 2012. http://www.bd.bibl.ita.br/tde_busca/arquivo.php?codArquivo=3218.

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Turbo-machinery design is clearly about making many decisions often under uncertainty and with multiple conflicting objectives. In this work, a computational tool has been developed to assist in the preliminary design optimization of supersonic turbines with application on turbopump rocket engines. It was proposed an evolutive approach based on genetic algorithm to automatize the process of selecting values, based on hands-on experience, for decision variables and fulfill simultaneously decisive compromises faced by the designer. At design point, exploring multiple optima solutions, the tool allows to fast estimate in a robust and accurate manner, performance, main dimensions, mass of rotor-wheel and the lower possible flow rate of the turbine. Out of this point, performance maps can be calculated varying rotational speed and pressure ratio. Because, it is involved by many objectives, a Pareto-optimum set is found. The search ends when the relation power-to-weight converges. The power-to-weight ratio characterizes a good option to relate performance and weight among distinct turbopump turbines. The empirical loss models adopted as well as the recommended values in the selection of decision variables were obtained in the Russian literature. To demonstrate the functionalities of tool, the single-stage of the Soviet RD109 rocket turbine was redesigned. The results were validated from those found in the engine';s atlas of construction and also, by a statistical method of calculation defined from an experimental study that estimates the maximum turbine efficiency.
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Ebaid, Munzer Shehadeh Yousef. "Design and construction of a small gas turbine to drive a permanent magnet high speed generator." Thesis, University of Hertfordshire, 2002. http://hdl.handle.net/2299/14046.

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Radial gas turbines engines have established prominence in the field of small turbomachinery because of their simplicity, relatively high performance and installation features. Thus they have been used in a variety of applications such as generator sets, small auxiliary power units (APu), air conditioning of aircraft cabins and hybrid electric vehicles turbines. The current research describes the design, manufacturing, construction and testing a radial type small gas turbine. The aim was to design and build the engine to drive directly a high-speed permanent magnet alternator running at 60000 rpmand developing a maximum of 60 W. This direct coupling arrangement produces a portable, light, compact, reliable and environment friendly power generator. These features make the generator set very attractive to use in many applications including emergency power generation for hospitals, in areas of natural disasters such as floods and earthquakes, in remote areas that cannot be served from the national grid, oil rigs, and in confined places of limited spaces. It is important to recognize that the design of the main components, that is, the inward flow radial UFR turbines, the centrifugal compressor and the combustion chamber involve consideration of aero-dynamics, thermodynamics, fluid mechanics, stress analysis, vibration analysis, selection of bearings, selection of suitable materials and the requirements for manufacturing. These considerations are all inter-linked and a procedure has been followed to reach an optimum design. This research was divided into three phases: phase I dealt with the complete design of the inward radial turbine, the centrifugal compressor, the power transmission shaft, the selection of combustion chamber and the bearing housing including the selection of bearings. Phase 2 dealt with mechanical consideration of the rotating components that is stress, thermal and vibration analyses of the turbine rotor, the impeller and the rotating shaft, respectively. Also it dealt with the selection of a suitable fuel and oil lubrication systems and a suitable starting system. Phase 3 dealt with the manufacturing of the gas turbine components, balancing the rotating components, assembling the engine and finally commissioning and then testing the engine. The current work in this thesis has put the light on a new design methodology on determining the optimum principal dimensions of the rotor and the impeller. This method, also, has defined the optimum number of blades and the axial length of the rotor and the impeller. Mathematical models linking the performance parameters and the design variables for the turbine and the compressor have been developed to assist in carrying out parametric studies to study the influence of the design parameters on the performance and on each other. Also, a new graphical matching procedure has been developed for the gas turbine components. This technique can serve as a valuable tool to determine the operating range and the engine running line. Furthermore, it would decide whether the gas turbine engine operates in a region of satisfactory compressor and turbine efficiencies.
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Tam, Kuok San. "Design and analysis of an electro-hydro-mechanical variable valve actuator for four-stroke automobile engines." Thesis, University of Macau, 2011. http://umaclib3.umac.mo/record=b2493685.

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Josselyn, Scott B. "Optimization of low thrust trajectories with terminal aerocapture." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2003. http://library.nps.navy.mil/uhtbin/hyperion-image/03Jun%5FJosselyn.pdf.

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Thesis (Aeronautical and Astronautical Engineer)--Naval Postgraduate School, June 2003.
Thesis advisor(s): I. Michael Ross, Steve Matousek. Includes bibliographical references (p. 149-150). Also available online.
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Books on the topic "Rocket engines – Design and construction"

1

Hindley, Keith B. Handbook of Russian rocket engines. York: Technology Detail, 1997.

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Bolonkin, Alexander. The development of Soviet rocket engines (for strategic missiles). Falls Church, VA: Delphic Associates, 1991.

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Grishin, S. D. Proektirovanie kosmicheskikh apparatov s dvigateli͡a︡mi maloĭ ti͡a︡gi. Moskva: Mashinostroenie, 1990.

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Belov, B. L. Razvitie teorii raketnykh dvigateleĭ: Do 40-kh godov XX stoletii︠a︡. Moskva: Institut istorii estestvoznanii︠a︡ i tekhniki RAN (IIET), 2006.

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Ye ti huo jian fa dong ji tui li shi she ji: Design for Thrust Chamber of Liquid Propellant Rocket Engines. Beijing Shi: Guo fang gong ye chu ban she, 2014.

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I︠A︡godnikov, D. A. Raketno-kosmicheskie dvigatelʹnye ustanovki: Materialy Vserossiĭskoĭ nauchno-tekhnicheskoĭ konferent︠s︡ii, Moskva, okti︠a︡brʹ 2013. Moskva: IIU MGOU, 2013.

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V, Glushko A., ed. Valentin Glushko: Konstruktor raketnykh dvigateleĭ i kosmicheskikh sistem. Sankt-Peterburg: Politekhnika, 2008.

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Ostapenko, I︠U︡ A. Raketoĭ sverknuvshai︠a︡ zhiznʹ: Stranit︠s︡y zhizni vydai︠u︡shchegosi︠a︡ konstruktora aviat︠s︡ionnoĭ i raketnoĭ tekhniki Aleksandra I︠A︡kovlevicha Berezni︠a︡ka. Dubna: Feniks+, 2012.

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Ye ti huo jian fa dong ji xian dai gong cheng she ji. Beijing shi: Zhongguo yu hang chu ban she, 2004.

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Ye ti huo jian fa dong ji xian dai gong cheng she ji. Beijing shi: Zhongguo yu hang chu ban she, 2004.

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Book chapters on the topic "Rocket engines – Design and construction"

1

Straub, D. "Design Criteria for Rocket Engines." In Thermofluiddynamics of Optimized Rocket Propulsions, 51–62. Basel: Birkhäuser Basel, 1989. http://dx.doi.org/10.1007/978-3-0348-9150-9_5.

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Pastrone, Dario, and Lorenzo Casalino. "Optimal Robust Design of Hybrid Rocket Engines." In Springer Optimization and Its Applications, 269–85. Cham: Springer International Publishing, 2016. http://dx.doi.org/10.1007/978-3-319-41508-6_10.

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Morgenweck, Daniel, Jutta Pieringer, and Thomas Sattelmayer. "Numerical Determination of Nozzle Admittances in Rocket Engines." In Notes on Numerical Fluid Mechanics and Multidisciplinary Design, 579–86. Berlin, Heidelberg: Springer Berlin Heidelberg, 2010. http://dx.doi.org/10.1007/978-3-642-14243-7_71.

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Traxinger, Christoph, Julian Zips, Christian Stemmer, and Michael Pfitzner. "Numerical Investigation of Injection, Mixing and Combustion in Rocket Engines Under High-Pressure Conditions." In Notes on Numerical Fluid Mechanics and Multidisciplinary Design, 209–21. Cham: Springer International Publishing, 2020. http://dx.doi.org/10.1007/978-3-030-53847-7_13.

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Abstract The design and development of future rocket engines severely relies on accurate, efficient and robust numerical tools. Large-Eddy Simulation in combination with high-fidelity thermodynamics and combustion models is a promising candidate for the accurate prediction of the flow field and the investigation and understanding of the on-going processes during mixing and combustion. In the present work, a numerical framework is presented capable of predicting real-gas behavior and nonadiabatic combustion under conditions typically encountered in liquid rocket engines. Results of Large-Eddy Simulations are compared to experimental investigations. Overall, a good agreement is found making the introduced numerical tool suitable for the high-fidelity investigation of high-pressure mixing and combustion.
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Fiedler, Torben, Joachim Rösler, Martin Bäker, Felix Hötte, Christoph von Sethe, Dennis Daub, Matthias Haupt, Oskar J. Haidn, Burkard Esser, and Ali Gülhan. "Mechanical Integrity of Thermal Barrier Coatings: Coating Development and Micromechanics." In Notes on Numerical Fluid Mechanics and Multidisciplinary Design, 295–307. Cham: Springer International Publishing, 2020. http://dx.doi.org/10.1007/978-3-030-53847-7_19.

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Abstract To protect the copper liners of liquid-fuel rocket combustion chambers, a thermal barrier coating can be applied. Previously, a new metallic coating system was developed, consisting of a NiCuCrAl bond-coat and a Rene 80 top-coat, applied with high velocity oxyfuel spray (HVOF). The coatings are tested in laser cycling experiments to develop a detailed failure model, and critical loads for coating failure were defined. In this work, a coating system is designed for a generic engine to demonstrate the benefits of TBCs in rocket engines, and the mechanical loads and possible coating failure are analysed. Finally, the coatings are tested in a hypersonic wind tunnel with surface temperatures of 1350 K and above, where no coating failure was observed. Furthermore, cyclic experiments with a subscale combustion chamber were carried out. With a diffusion heat treatment, no large-scale coating delamination was observed, but the coating cracked vertically due to large cooling-induced stresses. These cracks are inevitable in rocket engines due to the very large thermal-strain differences between hot coating and cooled substrate. It is supposed that the cracks can be tolerated in rocket-engine application.
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Manikandan, Kabaleeswaran, Boyapati Krishna Vamsi, Pola Anusha, and Chaturya Reddy. "Analysis of Internal Hydrodynamic Behavior of Pressure Swirl Atomizer for Liquid Rocket Engines." In Advances in Design and Thermal Systems, 261–70. Singapore: Springer Singapore, 2021. http://dx.doi.org/10.1007/978-981-33-6428-8_20.

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Armbruster, Wolfgang, Justin S. Hardi, and Michael Oschwald. "Experimental Investigation of Injection-Coupled High-Frequency Combustion Instabilities." In Notes on Numerical Fluid Mechanics and Multidisciplinary Design, 249–62. Cham: Springer International Publishing, 2020. http://dx.doi.org/10.1007/978-3-030-53847-7_16.

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Abstract Self-excited high-frequency combustion instabilities were investigated in a 42-injector cryogenic rocket combustor under representative conditions. In previous research it was found that the instabilities are connected to acoustic resonance of the shear-coaxial injectors. In order to gain a better understanding of the flame dynamics during instabilities, an optical access window was realised in the research combustor. This allowed 2D visualisation of supercritical flame response to acoustics under conditions similar to those found in European launcher engines. Through the window, high-speed imaging of the flame was conducted. Dynamic Mode Decomposition was applied to analyse the flame dynamics at specific frequencies, and was able to isolate the flame response to injector or combustion chamber acoustic modes. The flame response at the eigenfrequencies of the oxygen injectors showed symmetric and longitudinal wave-like structures on the dense oxygen core. With the gained understanding of the BKD coupling mechanism it was possible to derive LOX injector geometry changes in order to reduce the risks of injection-coupled instabilities for future cryogenic rocket engines.
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Perakis, Nikolaos, and Oskar J. Haidn. "Experimental and Numerical Investigation of CH$$_4$$/O$$_2$$ Rocket Combustors." In Notes on Numerical Fluid Mechanics and Multidisciplinary Design, 359–79. Cham: Springer International Publishing, 2020. http://dx.doi.org/10.1007/978-3-030-53847-7_23.

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Abstract The experimental investigation of sub-scale rocket engines gives significant information about the combustion dynamics and wall heat transfer phenomena occurring in full-scale hardware. At the same time, the performed experiments serve as validation test cases for numerical CFD models and for that reason it is vital to obtain accurate experimental data. In the present work, an inverse method is developed able to accurately predict the axial and circumferential heat flux distribution in CH$$_4$$/O$$_2$$ rocket combustors. The obtained profiles are used to deduce information about the injector-injector and injector-flame interactions. Using a 3D CFD simulation of the combustion and heat transfer within a multi-element thrust chamber, the physical phenomena behind the measured heat flux profiles can be inferred. A very good qualitative and quantitative agreement between the experimental measurements and the numerical simulations is achieved.
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Decher, Reiner. "More Components: Inlets, Mixers, and Nozzles." In The Vortex and The Jet, 137–54. Singapore: Springer Singapore, 2022. http://dx.doi.org/10.1007/978-981-16-8028-1_13.

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AbstractTheintegrationof a gas turbine engine into a functioning jet propulsion engine for an airplane requires more components: inlets and nozzles. For the inlet, the special care exercised to avoid ingestion of boundary layers air is described. The design features of nozzles are described and extended to include discussion of more extreme configurations such as those found on rocket engines.
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Akhmedzyanov, D. A., and A. E. Kishalov. "Computer-Aided Design and Construction Development of the Main Elements of Aviation Engines." In Lecture Notes in Electrical Engineering, 693–702. Cham: Springer International Publishing, 2020. http://dx.doi.org/10.1007/978-3-030-39225-3_76.

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Conference papers on the topic "Rocket engines – Design and construction"

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Kobald, M., C. Schmierer, U. Fischer, K. Tomilin, A. Petrarolo, and M. Rehberger. "The HyEnD stern hybrid sounding rocket project." In Progress in Propulsion Physics – Volume 11. Les Ulis, France: EDP Sciences, 2019. http://dx.doi.org/10.1051/eucass/201911025.

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The student team Hybrid Engine Development (HyEnD) of the University of Stuttgart is taking part with the Institute of Space Systems (IRS) in the DLR educational program STERN (Studentische Experimentalraketen). This program supports students at German universities to design, build, and launch an experimental rocket within a 3-year project time frame. HyEnD is developing a hybrid rocket called HEROS (Hybrid Experimental Rocket Stuttgart) with a design thrust of 10 kN, a total impulse of over 100 kN·s, and an expected liftoff weight up to 175 kg. HEROS is planned to be launched in October 2015 from Esrange in Sweden to an expected flight altitude of 40 to 50 km. The current altitude record for amateur rockets in Europe is at approximately 21 km. The propulsion system of HEROS is called HyRES (Hybrid Rocket Engine Stuttgart) and uses a paraffin-based solid fuel and nitrous oxide (N2O) as a liquid oxidizer. The development and the test campaign of HyRES is described in detail. The main goals of the test campaign are to achieve a combustion efficiency higher than 90% and provide stable operation with low combustion chamber pressure fluctuations. The successful design and testing of the HyRES engine was enabled by the evaluation and characterization of a small-scale demonstrator engine. The 500-newton hybrid rocket engine, called MIRAS (MIcro RAkete Stuttgart), has also been developed in the course of the STERN project as a technology demonstrator. During this test campaign, a ballistic characterization of paraffin-based hybrid rocket fuels with different additives in combination with N2O and a performance evaluation were carried out. A wide range of operating conditions, fuel compositions, injector geometries, and engine configurations were evaluated with this engine. Effects of different injector geometries and postcombustion chamber designs on the engine performance were analyzed. Additionally, the appearance of combustion instabilities under certain conditions, their effects, and possible mitigation techniques were also investigated. Concluding, the development and construction of an advanced, lightweight hybrid sounding rocket for the given requirements and budget within the DLR STERN program are described herein. The most important parts include a high thrust hybrid rocket engine, the development of a light weight oxidizer tank, pyrotechnical valves, carbon fiber rocket structure, recovery systems, and onboard electronics.
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SAGER, PAUL. "Design considerations in clustering nuclear rocket engines." In 28th Joint Propulsion Conference and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1992. http://dx.doi.org/10.2514/6.1992-3587.

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GUNN, STAN. "Design of second-generation nuclear thermal rocket engines." In 26th Joint Propulsion Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1990. http://dx.doi.org/10.2514/6.1990-1954.

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Smith, Nathan. "Facility design and testing of micro-hybrid rocket engines." In 39th Aerospace Sciences Meeting and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2001. http://dx.doi.org/10.2514/6.2001-5.

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Casalino, Lorenzo, Filippo Masseni, and Dario Pastrone. "Comparison of Robust Design Approaches for Hybrid Rocket Engines." In 53rd AIAA/SAE/ASEE Joint Propulsion Conference. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2017. http://dx.doi.org/10.2514/6.2017-4642.

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HUYNH, C., A. GHAFOURIAN, S. MAHALINGAM, and J. DAILY. "Combustor design for atomization study in liquid rocket engines." In 30th Aerospace Sciences Meeting and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1992. http://dx.doi.org/10.2514/6.1992-465.

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Casalino, Lorenzo, Filippo Masseni, and Dario Pastrone. "Uncertainty Analysis and Robust Design for Hybrid Rocket Engines." In 2018 Joint Propulsion Conference. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2018. http://dx.doi.org/10.2514/6.2018-4839.

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Raghavan, Madhusudan. "The Variable Rocker-Arm Mechanism." In ASME 2005 International Design Engineering Technical Conferences and Computers and Information in Engineering Conference. ASMEDC, 2005. http://dx.doi.org/10.1115/detc2005-85150.

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We present an investigation of a novel variable valve-lift mechanism known as the Variable Rocker-Arm mechanism (VR-Arm, for short). This mechanism has simple construction, low friction due to rolling contact, and provides continuously variable lift. Dynamic analysis of the mechanism using ADAMS shows that frictional losses at the intake camshaft are of the order of 1 Newton-meter for a four-cylinder engine.
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MEYER, MICHAEL. "Design issues for lunar in situ aluminum/oxygen propellant rocket engines." In Aerospace Design Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1992. http://dx.doi.org/10.2514/6.1992-1185.

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LIMERICK, C. "Component design concerns for deep throttling H2/O2 rocket engines." In 27th Joint Propulsion Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1991. http://dx.doi.org/10.2514/6.1991-2209.

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Reports on the topic "Rocket engines – Design and construction"

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Morais, Carla, António Coelho, Alexandre Jacinto, and Marta Varzim, eds. The I SEA Project: Digital Publications. Faculdade de Ciências da Universidade do Porto, October 2020. http://dx.doi.org/10.24840/2020/978-989-746-279-5.

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The I SEA project aimed at the development of a non-obtrusive, valid and replicable method to evaluate audience attitudes about science communication projects through an immersive virtual reality environment that can improve exhibitions while educating and empowering citizens. To achieve the objectives of this highly complex, highly interdisciplinary, and innovative project, a permanent articulation of the scientific approach with the technical and design development took place, aiming the construction of the non- invasive evaluation method. Because it is an intricate project, it required constant iterations and interactions among the team members. So, we’ve learned somehow to consider limitations as engines for developing the project, instead of seeing them as obstacles.
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