Journal articles: '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.

Full text

Abstract:

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.

APA, Harvard, Vancouver, ISO, and other styles

2

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.

Full text

Abstract:

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.

APA, Harvard, Vancouver, ISO, and other styles

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.

Full text

Abstract:

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.

APA, Harvard, Vancouver, ISO, and other styles

4

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.

Full text

Abstract:

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

APA, Harvard, Vancouver, ISO, and other styles

5

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.

Full text

Abstract:

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.

APA, Harvard, Vancouver, ISO, and other styles

6

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.

Full text

APA, Harvard, Vancouver, ISO, and other styles

7

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.

Full text

Abstract:

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.

APA, Harvard, Vancouver, ISO, and other styles

8

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.

Full text

Abstract:

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.

APA, Harvard, Vancouver, ISO, and other styles

9

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.

Full text

Abstract:

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.

APA, Harvard, Vancouver, ISO, and other styles

10

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.

Full text

Abstract:

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: BrÃ¼ker 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.

APA, Harvard, Vancouver, ISO, and other styles

11

Szwaja, Stanisław, and Mirosław Szymkowiak. "The Szymkowiak’s over-expanded cycle in the rocker engine with the variable compression ratio – kinematics." Combustion Engines 189, no. 2 (October 30, 2021): 68–72. http://dx.doi.org/10.19206/ce-143157.

Full text

Abstract:

The article discusses the innovative concept of the over-expanded thermodynamic cycle, the author of which is the Polish engineer-designer Mirosław Szymkowiak. This cycle is realized on the basis of a new and innovative, previously unknown design, of a piston-crankshaft linkage mechanism with the aid of an additional element known as a rocker arm. Additionally, the proposed mechanism allows for a smooth change of the compression/expansion ratio of the engine during its operation. In the beginning, the earlier conceptions of the rocker engine developed by Szymkowiak were presented, and then the main construction assumptions and kinematic calculations were described. It was confirmed, that the developed linkage has big potential in improving the engine's thermal efficiency by approximately 12% relative. Additionally, it significantly reduces the exhaust gas pressure, when the exhaust valve is opened, therefore, contributes to the reduction of the noise emitted by the engine.

APA, Harvard, Vancouver, ISO, and other styles

12

Салич, В. Л. "A numerical study of the working process in the chamber of a thruster rocket engine based on oxygen-hydrogen fuel." Numerical Methods and Programming (Vychislitel'nye Metody i Programmirovanie), no. 2 (June 30, 2015): 187–95. http://dx.doi.org/10.26089/nummet.v16r219.

Full text

Abstract:

Применение численного моделирования смесеобразования и горения в процессе проектирования камеры кислородно-водородного ракетного двигателя малой тяги позволило в короткие сроки получить конструкцию, обеспечивающую высокие энергетические характеристики, что впоследствии было подтверждено экспериментально. В настоящей статье представлены результаты расчетов, полученные при использовании различных моделей турбулентности (моделей на основе гипотезы турбулентной вязкости и рейнольдсовых напряжений) и моделей химического взаимодействия (моделей тонкого фронта пламени и диссипации вихря), а также результаты, полученные на различных типах и размерностях расчетной сетки. В результате исследований установлено, что тип сетки не оказывает существенного влияния на результаты моделирования; предпочтения отданы модели рейнольдсовых напряжений и модели диссипации вихря. Приводятся сопоставления характеристик камеры на различных режимах работы, полученных экспериментально и по результатам моделирования. The application of numerical simulation of mixture formation and combustion during the design of the chamber of a thruster rocket engine based on oxygen-hydrogen fuel allows one to rapidly develop a construction with high-power characteristics, which was later confirmed experimentally. This paper considers the results of computations obtained by using the turbulence models based on the hypothesis of turbulent viscosity and Reynolds stresses and using the chemical interaction models (the models of a thin flame front and the eddy dissipation). The numerical results obtained on the basis of computational grids of various types and dimensions are discussed. It is shown that the grid type has little effect on the simulation results and that the Reynolds stress model and the eddy dissipation model are preferable. The characteristics of the chamber obtained experimentally and numerically for various modes of operation are compared.

APA, Harvard, Vancouver, ISO, and other styles

13

Partola, I. S. "Design of liquid-propellant rocket engines." Journal of Machinery Manufacture and Reliability 41, no. 6 (November 2012): 492–98. http://dx.doi.org/10.3103/s1052618812060118.

Full text

APA, Harvard, Vancouver, ISO, and other styles

14

Kamennov, P. B. "PRC’S MILITARY-INDUSTRIAL COMPLEX UNDER XI JINPING." Outlines of global transformations: politics, economics, law 10, no. 5 (December 20, 2017): 135–51. http://dx.doi.org/10.23932/2542-0240-2017-10-5-135-151.

Full text

Abstract:

The article is dedicated to the achievements and problems of the Chinese military-industrial complex and its role in the economy building during the Xi Jinping` s term (2012–2017). The evolution of the problem is particularly interesting in the context of the realization of the National Security modernization Program of China 2050. The author provides a description of the current state of military-industrial complex fields. Chinese atomic industry has been growing rapidly during the last years. There are significant achievements in the field of the rocket and space industry. Aviation industry, and in particular the aviation engine construction, has been traditionally considered as underdeveloped in comparison with the American, Western European and Russian aviation. In this sense, the Chinese government endeavors to improve the situation. Microelectronics, which has also been rather underdeveloped, is expected to obtain a new quality. The shipbuilding industry in China nowadays gives positive examples of design and building heavy-tonnage vessels, such as nuclear-powered submarines. Apart from the military and technological achievements, many efforts are put in order to reduce the level of the Chinese dependence on the import of foreign technologies. The author stresses the significance of the military industry conversion practices, which used to be one of the main driving forces of the China`s economic growth in 1980–1990s. The so-called military civil integration is one of the most effective strategies aimed at overcoming the technological backwardness of the military- industrial complex. It contributes to the technological growth in the fields of security due to the practices of redirection the civil technologies into the military field. Such an approach allows China to avoid the Western military embargo.

APA, Harvard, Vancouver, ISO, and other styles

15

Ryzhkov, V. V. "The structure of knowledge base to support the development of low-thrust liquid-propellant rocket engines based on computer technologies." VESTNIK of Samara University. Aerospace and Mechanical Engineering 20, no. 4 (January 19, 2022): 28–39. http://dx.doi.org/10.18287/2541-7533-2021-20-4-28-39.

Full text

Abstract:

Some data on the knowledge base to support the development of low-thrust liquid-propellant rocket engines using computer technologies are presented. The structure of the base is proposed on the basis of characteristic features of engines, including the purpose, fuel components, physical principles of organizing the work process of the engines, etc. The presence of electronic versions of schematic diagrams, configuration and the main achieved characteristics in the database will make it possible to choose effective design solutions at the design stage of new products. In the future these solutions will lead to the required parameters and characteristics of the low-thrust rocket engines being developed. The description of the engine used in the database allows assessing the capabilities of the engineering solution used in the design, as well as tracing the development trends of a particular direction in rocket propulsion. The peculiarity of the base being created is that, in parallel with the information about low-thrust rocket engines, the data on their components and accessories is accumulating, which can also be used in new developments. Given the growing volume of the knowledge base on low-thrust rocket engines, some forms of communications are presented that make it possible to quickly find the required information, but requires certain ordering of the design data already at the initial stages.

APA, Harvard, Vancouver, ISO, and other styles

16

Shulga, V. A., and A. V. Dibrivny. "Design Office of Liquid Rocket Engines is 60." Kosmičeskaâ tehnika. Raketnoe vooruženie 2018, no. 2 (September 14, 2018): 3–7. http://dx.doi.org/10.33136/stma2018.02.003.

Full text

APA, Harvard, Vancouver, ISO, and other styles

17

Ray, A., M. S. Holmes, and C. F. Lorenzo. "Life extending controller design for reusable rocket engines." Aeronautical Journal 105, no. 1048 (June 2001): 315–22. http://dx.doi.org/10.1017/s0001924000012197.

Full text

Abstract:

Abstract The goal of life extending control (LEC) is to enhance structural durability of complex mechanical systems, such as aircraft, spacecraft, and energy conversion devices, without incurring any significant loss of performance. This paper presents a concept of robust life-extending controller design for reusable rocket engines, similar to the Space Shuttle Main Engine (SSME), via damage mitigation in both fuel and oxidiser turbines while achieving the required performance for transient responses of the main combustion chamber pressure and the oxidant/fuel mixture ratio. The design procedure makes use of a combination of linear robust control synthesis and nonlinear optimisation techniques. Results of simulation experiments on the model of a reusable rocket engine are presented to this effect.

APA, Harvard, Vancouver, ISO, and other styles

18

Etele, J., T. Waung, and D. J. Cerantola. "Exchange Inlet Design for Rocket-Based Combined-Cycle Engines." Journal of Propulsion and Power 28, no. 5 (September 2012): 1026–36. http://dx.doi.org/10.2514/1.b34531.

Full text

APA, Harvard, Vancouver, ISO, and other styles

19

Lee, Dae-Sung, Chang-Ho Choi, Jin-Han Kim, and Soo-Seok Yang. "Design of a Turbine System for Liquid Rocket Engines." Journal of Fluid Machinery 5, no. 4 (December 1, 2002): 11–18. http://dx.doi.org/10.5293/kfma.2002.5.4.011.

Full text

APA, Harvard, Vancouver, ISO, and other styles

20

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.

Full text

Abstract:

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.

APA, Harvard, Vancouver, ISO, and other styles

21

Siva Vasanth, A., J. Stalin, and A. Benzigar Rajan. "Instrumentation design considerations for automatic ground testing of rocket engines." Indian Journal of Cryogenics 46, no. 1 (2021): 171–77. http://dx.doi.org/10.5958/2349-2120.2021.00030.3.

Full text

APA, Harvard, Vancouver, ISO, and other styles

22

Santana Jr., A., M. S. Silva, P. T. Lacava, and L. C. S. Góes. "ACOUSTIC CAVITIES DESIGN PROCEDURES." Revista de Engenharia Térmica 6, no. 2 (December 31, 2007): 27. http://dx.doi.org/10.5380/reterm.v6i2.61687.

Full text

Abstract:

Combustion instability is recognized as one of the major problems frequently faced by engineers during the development of either liquid or solid propellant rocket engines. The performance of the engine can be highly affected by these high frequencies instabilities, possibly leading the rocket to an explosion. The main goal while studying combustion chambers instability, either by means of baffles or acoustic absorbers, is to achieve the stability needed using the simplest possible manner. This paper has the purpose of studying combustion chambers instabilities, as well as the design of acoustic absorbers capable of reducing their eigenfrequencies. Damping systems act on the chamber eigenfrequency, which has to be, therefore, previously known.

APA, Harvard, Vancouver, ISO, and other styles

23

Guo, Feng, Wenguo Luo, Feng Gui, Jianfeng Zhu, Yancheng You, and Fei Xing. "Efficiency Analysis and Integrated Design of Rocket-Augmented Turbine-Based Combined Cycle Engines with Trajectory Optimization." Energies 13, no. 11 (June 5, 2020): 2911. http://dx.doi.org/10.3390/en13112911.

Full text

Abstract:

An integrated analysis method for a rocket-augmented turbine-based combined cycle (TBCC) engine is proposed based on the trajectory optimization method of the Gauss pseudospectral. The efficiency and energy of the vehicles with and without the rocket are analyzed. Introducing an appropriate rocket to assist the TBCC-powered vehicle will reduce the total energy consumption of drag, and increase the vehicle efficiency in the transonic and the mode transition. It results in an increase in the total efficiency despite a reduction in engine efficiency. Therefore, introducing a rocket as the auxiliary power is not only a practical solution to enable flight over a wide-speed range when the TBCC is incapable but also probably an economical scheme when the the TBCC meets the requirements of thrust. When the vehicle drag is low, the rocket works for a short time and its optimal relative thrust is small. Thus, the TBCC combined with a booster rocket will be a more simple and suitable scheme. When the vehicle drag is high, the operating time of the rocket is long and the optimal relative thrust is large. The specific impulse has a significant impact on the flight time and the total fuel consumption. Accordingly, the combination form for the rocket-based combined cycle (RBCC) engines and the turbine will be more appropriate to obtain higher economic performance.

APA, Harvard, Vancouver, ISO, and other styles

24

Casalino, Lorenzo, Andrea Ferrero, Filippo Masseni, and Dario Pastrone. "Emission-Driven Hybrid Rocket Engine Optimization for Small Launchers." Aerospace 9, no. 12 (December 9, 2022): 807. http://dx.doi.org/10.3390/aerospace9120807.

Full text

Abstract:

Hybrid rocket engines are a green alternative to solid rocket motors and may represent a low-cost alternative to kerosene fueled rockets, while granting performance and control features similar to that of typical storable liquid rocket engines. In this work, the design of a three-stage hybrid launcher is optimized by means of a coupled procedure: an evolutionary algorithm optimizes the engine design, whereas an indirect optimization method optimizes the corresponding ascent trajectory. The trajectory integration also provides the vertical emission profiles required for the evaluation of the environmental impact of the launch. The propellants are a paraffin-based wax and liquid oxygen. The vehicle is launched from the ground and uses an electric turbo pump feed system. The initial mass is given (5000 kg) and the insertion of the payload into a 600-km circular, and polar orbit is considered as a reference mission. Clusters of similar hybrid rocket engines, with only few differences, are employed in all stages to reduce the development and operational costs of the launcher. Optimization is carried out with the aim of maximizing the payload mass and then minimizing the overall environmental impact of the launch. The results show that satisfactory performance is achievable also considering rocket polluting emissions: the carbon footprint of the launch can be reduced by one fourth at the cost of a 5-kg payload mass reduction.

APA, Harvard, Vancouver, ISO, and other styles

25

Glebov, G. A., and S. A. Vysotskaya. "On the question of solid-propellant rocket engine design preventing unstable operation in the combustion chamber." Journal of «Almaz – Antey» Air and Space Defence Corporation, no. 4 (December 30, 2017): 63–72. http://dx.doi.org/10.38013/2542-0542-2017-4-63-72.

Full text

Abstract:

The paper presents results of a numerical investigation concerning the effect that the flow duct shape and combustion rate equation have on the gas dynamic vortex flow pattern and self-excited pressure oscillations in the combustion chamber of a solid-propellant rocket engine. We provide guidelines on upgrading solid-propellant rocket engines in order to decrease the magnitude of pressure pulses in the case of pulsating combustion.

APA, Harvard, Vancouver, ISO, and other styles

26

Son, Min, Kanmaniraja Radhakrishnan, Jaye Koo, Oh Chae Kwon, and Heuy Dong Kim. "Design Procedure of a Movable Pintle Injector for Liquid Rocket Engines." Journal of Propulsion and Power 33, no. 4 (July 2017): 858–69. http://dx.doi.org/10.2514/1.b36301.

Full text

APA, Harvard, Vancouver, ISO, and other styles

27

Konokh, V. I., V. S. Boiko, A. B. Troyak, and A. V. Ivashura. "Electromagnetic Valves Developed by Yuzhnoye SDO Liquid Rocket Engines Design Office." Kosmičeskaâ tehnika. Raketnoe vooruženie 2018, no. 2 (September 14, 2018): 34–48. http://dx.doi.org/10.33136/stma2018.02.034.

Full text

APA, Harvard, Vancouver, ISO, and other styles

28

KANDA, Takeshi, Yohei OGAWA, Daizo SUGIMORI, and Makoto KOJIMA. "Conceptual Design Model of High-Altitude Test Stand for Rocket Engines." TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 59, no. 3 (2016): 161–69. http://dx.doi.org/10.2322/tjsass.59.161.

Full text

APA, Harvard, Vancouver, ISO, and other styles

29

Ramesh, Davood, Hasan Karimi M., and Massoud Shahheidari. "Cycle optimization of the staged combustion rocket engines." Aircraft Engineering and Aerospace Technology 89, no. 2 (March 6, 2017): 304–13. http://dx.doi.org/10.1108/aeat-12-2013-0229.

Full text

Abstract:

Purpose The purpose of this paper is to introduce new and modified “staged combustion” cycles in the form of engineering algorithm as a possible propulsion contender for future aerospace vehicle to achieve the highest possible “total impulse” to “mass” of propulsion system. Design/methodology/approach In this regard, the mathematical cycle model is formed to calculate the engine’s parameters. In addition, flow conditions (pressure, temperature, flow rate, etc). in the chamber, nozzle and turbopump are assessed based on the results of turbo machinery power balance and initial data such as thrust, propellant mixture ratio and specifications. The developed code has been written in the modern, object-oriented C++ programming language. Findings The results of the developed code are compared with the Russian RD180 engine which demonstrates the superiority and capability of new “thermodynamic diagrams”. Research limitations/implications This algorithm is under constraint to control the critical variation of combustion pressure, turbine rpm, pump cavitation and turbine temperature. It is imperative to emphasize that this paper is limited to “oxidizer-rich staged combustion” engines with “single pre-burner”. Originality/value This study sheds light on using fuel booster turbopump and the second-stage fuel pump to moderate the effect of cavitation on pumps which reduces tank pressure and, as a consequence, decreases the propulsion system weight.

APA, Harvard, Vancouver, ISO, and other styles

30

Liu, Chao Ying, and Ran Duan. ""A+T·CDIO"-Architecture Major Oriented Talents Cultivation Objectives Establishment and Curriculum System Construction - Using Architecture Major of Ningbo University of Technology as Example." Applied Mechanics and Materials 357-360 (August 2013): 2806–9. http://dx.doi.org/10.4028/www.scientific.net/amm.357-360.2806.

Full text

Abstract:

How to transform A+T·CDIO engineering education teaching model of architecture major to executable curriculum system? Approach is job-adapting oriented, aiming to adapt social and industrial environment, using A+T·CDIO curriculum system as carrier. This article describes talents cultivation objectives featured as A 380 and their three construction basis for architecture major, and A+T·CDIO curriculum system featured as eight engines and three-section rocket and their four construction principles.

APA, Harvard, Vancouver, ISO, and other styles

31

Anoop, C. R., John Bejoy, Thomas Tharian, and P. V. Venkitakrishnan. "Evaluation of Mechanical Properties of the Rocket Thrust Chamber Material under Flight Conditions." Materials Science Forum 830-831 (September 2015): 219–22. http://dx.doi.org/10.4028/www.scientific.net/msf.830-831.219.

Full text

Abstract:

Thrust chamber of rocket engines often operate under conditions of rapid heating environments with temperatures approaching the melting points of the materials involved. Therefore, for the optimum design of rocket engines, it is necessary to obtain the properties of thrust chamber materials under actual operating conditions. The heating rates and loading rates in conventional (laboratory) tensile stress-strain tests, which are intended to evaluate the high temperature tensile properties of the materials, are usually very low compared to that actually encountered in rocket engines. The heating rates in conventional stress-strain tests are of the order of 0.4K per second to 0.5K per second only, whereas the combustion and aerodynamic heating rates in rockets and re-entry vehicles will usually exceed 100K per second. In this context, a very important beginning has been made to experimentally determine the high temperature tensile properties of KC20WN (a cobalt based superalloy) used in earth storable liquid rocket engine thrust chamber under conditions of rapid rates of heating which actually exists during flight. These investigations have shown that the elevated temperature strength of KC20WN depends upon the heating rate (or heating time) and can be considerably higher for rapid-heating conditions than for conventional heating conditions.

APA, Harvard, Vancouver, ISO, and other styles

32

Kang, Sang Hun, and Jesun Jang. "Design and Hot Fire Tests of the Pyrostarter for Liquid Rocket Engines." Journal of the Korean Society of Propulsion Engineers 18, no. 3 (June 1, 2014): 48–55. http://dx.doi.org/10.6108/kspe.2014.18.3.048.

Full text

APA, Harvard, Vancouver, ISO, and other styles

33

Elkin, Andrey. "Rocket engines for spacecraft on fluidized solid propellants, their design and thermodynamics." Thermal processes in engineering 13, no. 11 (2021): 509–18. http://dx.doi.org/10.34759/tpt-2021-13-11-509-518.

Full text

APA, Harvard, Vancouver, ISO, and other styles

34

Ryzhkov, V. V., I. I. Morozov, and E. A. Lapshin. "Computer-aided design of low-thust rocket engines using the domain-specific knowledge database and CAE / CAD systems." VESTNIK of Samara University. Aerospace and Mechanical Engineering 18, no. 4 (January 21, 2020): 106–16. http://dx.doi.org/10.18287/2541-7533-2019-18-4-106-116.

Full text

Abstract:

The paper presents approaches to computer-aided design of low-thrust thrust rocket engines using an extensive knowledge base that allows making basic technical decisions that determine the conceptual design of the engine, based on the developed algorithm of this process. The procedure of creating an electronic 3D-model of a low-thrust rocket engine fueled by gaseous oxygen-hydrogen in the environment of the graphical complex UNIGRAPHICS is described. 3D electronic models of the main elements of a rocket engine with a thrust of P = 25 N were obtained, with subsequent virtual assembly of all components, including the components comprised in the knowledge base, providing the development, among other things, of design documentation, creation of a production environment based on an electronic engine model, preparation for the product manufacturing and the manufacturing proper.

APA, Harvard, Vancouver, ISO, and other styles

35

Messineo, Jérôme, and Toru Shimada. "Theoretical Investigation on Feedback Control of Hybrid Rocket Engines." Aerospace 6, no. 6 (June 3, 2019): 65. http://dx.doi.org/10.3390/aerospace6060065.

Full text

Abstract:

Despite the fact that hybrid propulsion offers significant benefits, it still suffers from some limitations such as the natural oxidizer to fuel ratio shift which induces variations of the engines’ performances while operating. To overcome that issue, Japan Aerospace Exploration Agency (JAXA) has been studying an innovative concept for several years based on the combination of controlled axial and radial oxidizer injections, called altering-intensity swirling-oxidizer-flow-type engine. This type of motor is theoretically capable of managing both the thrust and the oxidizer to fuel ratio independently and instantaneously by using a feedback control loop. To be effective, such engines would require in-flight instantaneous and precise thrust and an oxidizer to fuel ratio measurements as well as an adapted feedback control law. The purpose of this study is to investigate the effect of measurement errors on the engine control and to propose a regulation law suitable for these motors. Error propagation analysis and regulation law are developed from fundamental equations of hybrid motors and applied in a case where resistor-based sensors are used for fuel regression rate measurement. This study proves the theoretical feasibility of hybrid engines feedback control while providing some methods to design the engine and regression rate sensors depending on the mission requirements.

APA, Harvard, Vancouver, ISO, and other styles

36

Vaulin, S. D., and K. I. Khazhiakhmetov. "The State-of-the-Art and Prospects of Aerospike Engines." Proceedings of Higher Educational Institutions. Маchine Building, no. 10 (739) (October 2021): 74–83. http://dx.doi.org/10.18698/0536-1044-2021-10-74-83.

Full text

Abstract:

Currently, there is a worldwide trend of growing interest in projects aimed at reducing the cost of spacecraft launches. The search for solutions to this topical issue reveals new requirements for the rocket engines. However, existing rocket engines are incapable of fully meeting modern requirements. Consideration of new technical solutions indicates the prospects of using aerospike engines, which have the property of self-regulation and can operate with optimal flow expansion throughout the entire operation. This property allows this type of engine to be used as a propulsion system for single-stage return launcher. However, aerospike engines have not been sufficiently studied at the moment and haven’t found widespread use. Therefore it is necessary to summarize the existing knowledge about aerospike and the aerospike research has been performed. As a result mathematical models of the workflows were created, methods of designing and optimization of the contour were determined, and a number of design and technological solutions were found. However, the mathematical models verified by experimental data were not found.

APA, Harvard, Vancouver, ISO, and other styles

37

Lamb, Thomas, Alex M. Loudon, and Robert J. Van Slyke. "The Lockheed Space Shuttle Rocket Retrieval Ship." Marine Technology and SNAME News 23, no. 02 (April 1, 1986): 109–22. http://dx.doi.org/10.5957/mt1.1986.23.2.109.

Full text

Abstract:

This paper describes the design and construction of a 200-ft rocket retrieval ship for West Coast space shuttle operations. The ship, MV Independence, is designed to retrieve the reusable solid rocket booster casings from each launch of the space shuttle. Construction design and technical management of the project are outlined and a typical rocket retrieval mission is described. Updated information on builder's trials and mission performance is appended.

APA, Harvard, Vancouver, ISO, and other styles

38

Rugescu, Radu Dan, Florin Radu Bacaran, and Stefan Catalin Predoiu. "Starting Transient Experiments in the MEC-80 Rocket Engine Scaled Ignition Device." Applied Mechanics and Materials 656 (October 2014): 110–17. http://dx.doi.org/10.4028/www.scientific.net/amm.656.110.

Full text

Abstract:

One of the main sources of concern within the design and manufacturing of rocket engines is the starting device, due to the relatively low reliability of the ignition of the main combustion process in these engines. Not only is necessary for the igniter to produce a sound ignition, but the starting pressure transient should manifest a tightly controlled build-up, in a very definite time interval. A specific research and development work was done by the authors in order to reliably assess the requirements for the ignition device of the newly developed compound, solid-liquid experimental rocket engine MEC-80 of the ADDA team. Computer predictions and scaled experiments on the constant volume combustion into a modified calorimeter were performed with notable results that led to the optimal design of the ignition device of the motor. The test results are presented with emphasize on the predictability of the ignition delay.

APA, Harvard, Vancouver, ISO, and other styles

39

Kim, Hye In, Tae-Seong Roh, Hwanil Huh, and Hyoung Jin Lee. "Development of Ultra-Low Specific Speed Centrifugal Pumps Design Method for Small Liquid Rocket Engines." Aerospace 9, no. 9 (August 28, 2022): 477. http://dx.doi.org/10.3390/aerospace9090477.

Full text

Abstract:

With the growth of the satellite industry, the demand for a propulsion system for small launch vehicles and spacecraft has increased. Small liquid rocket engines may require Ultra-Low specific speed centrifugal pumps due to the low required thrust and volumetric flow rate and high combustion chamber pressure. Therefore, in this study, a design method of Ultra-Low specific speed centrifugal pumps for several hundred Newton class small liquid rocket engines was developed by combining various empirical formulas. In addition, centrifugal pump impellers were designed using the Stepanoff method, which is typically used in pump design, and the circular arc method. The most appropriate method for designing Ultra-Low specific speed centrifugal pumps was determined through a comparative analysis with other methods and validated through CFD. As a result, the pump designed using the proposed method exhibited a performance of pumping and suction superior to the Stepanoff method. Although the number of arcs did not considerably influence the pump performance, the single arc method was confirmed to be the most appropriate design approach in terms of the design productivity and simplicity.

APA, Harvard, Vancouver, ISO, and other styles

40

Khan, Suniya Sadullah, Ihtzaz Qamar, Muhammad Umer Sohail, Raees Fida Swati, Muhammad Azeem Ahmad, and Saad Riffat Qureshi. "Comparison of Optimization Techniques and Objective Functions Using Gas Generator and Staged Combustion LPRE Cycles." Applied Sciences 12, no. 20 (October 17, 2022): 10462. http://dx.doi.org/10.3390/app122010462.

Full text

Abstract:

This paper compares various optimization techniques and objective functions to obtain optimum rocket engine performances. This research proposes a modular optimization framework that provides an optimum design for Gas Generator (GG) and Staged Combustion (SC) Liquid Propellant Rocket Engines. This process calculates the ideal rocket engine performance by applying seven different optimization techniques: Simulated Annealing (SA), Nelder Mead (NM), Cuckoo Search Algorithm (CSA), Particle Swarm Optimization (PSO), Pigeon-Inspired Optimization (PIO), Genetic Algorithm (GA) and a novel hybrid GA-PSO technique named GA-Swarm. This new technique combines the superior search capability of GA with the efficient constraint matching capability of PSO. This research also compares objective functions to determine the most suitable function for GG and SC cycle rocket engines. Three single objective functions are used to minimize the Gross Lift-Off Weight and to maximize Specific Impulse and the Thrust-to-Weight ratio. A fourth multiobjective function is used to simultaneously maximize both Specific Impulse and Thrust-to-Weight ratio. This framework is validated against a pump-fed rocket, and results are within 1% of the actual rocket engine mass. The results of this research indicate that PSO and GA-Swarm produce optimum results for all objective functions. Finally, the most suitable objective function to use while comparing these two cycles is the Gross Lift-Off Weight.

APA, Harvard, Vancouver, ISO, and other styles

41

Razin, A. F. "The Problem of Optimum Design of Composite Housings of Solid Propellant Rocket Engines." Mechanics of Solids 53, no. 4 (July 2018): 418–26. http://dx.doi.org/10.3103/s0025654418040076.

Full text

APA, Harvard, Vancouver, ISO, and other styles

42

Ryzhkov, V., and E. Lapshin. "Computer Design Technology of the Small Thrust Rocket Engines Using CAE / CAD Systems." IOP Conference Series: Materials Science and Engineering 302 (January 2018): 012032. http://dx.doi.org/10.1088/1757-899x/302/1/012032.

Full text

APA, Harvard, Vancouver, ISO, and other styles

43

Frank, Christopher P., Olivia J. Pinon-Fischer, Dimitri N. Mavris, and Clémence M. Tyl. "Design Methodology for the Performance, Weight, and Economic Assessment of Chemical Rocket Engines." Journal of Aerospace Engineering 30, no. 1 (January 2017): 04016071. http://dx.doi.org/10.1061/(asce)as.1943-5525.0000668.

Full text

APA, Harvard, Vancouver, ISO, and other styles

44

Лавров, С. В., and Б. К. Терпогосова. "CHOOSING RATIONAL ENGINEERING SOLUTIONS AND DESIGN CRITERIA OF THROTTABLE SOLID FUEL ROCKET ENGINES." Южно-Сибирский научный вестник, no. 5(45) (October 31, 2022): 78–81. http://dx.doi.org/10.25699/sssb.2022.45.5.010.

Full text

Abstract:

В статье проанализированы основные сложности процесса выбора рациональных технических решений и параметров регулируемых двигательных установок на твердом топливе (ТРДУ) на начальном этапе проектирования. Выявлена актуальность повышения обоснованности процесса выбора, предложены наиболее значимые показатели для оценки качества множества альтернативных вариантов ТРДУ, выполнена постановка задачи оптимизации основных проектных параметров. The article analyzes the main difficulties of the process of choosing rational technical solutions and design criteria of throttable solid fuel rocket engines (TSRE) at the initial design stage. The relevance of increasing the validity of the selection process is revealed, the most significant indicators for assessing the quality of a variety of alternative options TSRE, the task of optimizing the main design criteria is formulated.

APA, Harvard, Vancouver, ISO, and other styles

45

Casalino, Lorenzo, Filippo Masseni, and Dario Pastrone. "Hybrid Rocket Engine Design Optimization at Politecnico di Torino: A Review." Aerospace 8, no. 8 (August 13, 2021): 226. http://dx.doi.org/10.3390/aerospace8080226.

Full text

Abstract:

Optimization of Hybrid Rocket Engines at Politecnico di Torino began in the 1990s. A comprehensive review of the related research activities carried out in the last three decades is here presented. After a brief introduction that retraces driving motivations and the most significant steps of the research path, the more relevant aspects of analysis, modeling and achieved results are illustrated. First, criteria for the propulsion system preliminary design choices (namely the propellant combination, the feed system and the grain design) are summarized and the engine modeling is presented. Then, the authors describe the in-house tools that have been developed and used for coupled trajectory and propulsion system design optimization. Both deterministic and robust-based approaches are presented. The applications that the authors analyzed over the years, starting from simpler hybrid powered sounding rocket to more complex multi-stage launchers, are then presented. Finally, authors’ conclusive remarks on the work done and their future perspective in the context of the optimization of hybrid rocket propulsion systems are reported.

APA, Harvard, Vancouver, ISO, and other styles

46

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.

Full text

Abstract:

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.

APA, Harvard, Vancouver, ISO, and other styles

47

Nae, Catalin, Irina-Carmen Andrei, Gabriela-Liliana Stroe, and Sorin Berbente. "Integration of Fuels Types and Chemical Properties with the Design of the Rocket Engine�s Bell Exhaust Nozzle and Combustion Chamber." Revista de Chimie 71, no. 1 (February 7, 2020): 436–44. http://dx.doi.org/10.37358/rc.20.1.7872.

Full text

Abstract:

The chemical properties of the fuels are crucial for obtaining the numerical accuracy during the design and performance analysis in case of liquid fuel propelled rocket engines, as well as the trajectory optimization. In this paper, the research was primarly focused on optimizing the numerical accuracy for non-linear two-dimensional approximation the Fuel Combustion Charts; secondarily, the investigation was carried on the design of the bell-nozzle of a liquid propelled rocket engine, taking into account the variation of the coefficients which are significant for expressing the fuels chemical properties. From the Fuel Combustion Charts, the authors selected a the LOX - Kerosene combination for propelling the rocket engine, due to the most convenient matching with the technology and material specifications, safety and environmental friendly requirements; from the LOX-Kerosene Charts, the authors have originally developed a method to obtain the expression of a non-linear approximation function of two variables. The design of the bell shaped nozzle and combustion chamber for a liquid propelled rocket engine was included, in purpose to illustrate the link between the LPRE design and the fuels types and chemical properties.

APA, Harvard, Vancouver, ISO, and other styles

48

Ponomarov, O. M., O. O. Dobrodomov, and O. V. Kulyk. "Prospects for the use of nitrogen-containing single-component rocket propellants." Technical mechanics 2022, no. 3 (October 3, 2022): 85–90. http://dx.doi.org/10.15407/itm2022.03.085.

Full text

Abstract:

The goal of this work is to analyze the possibility of using existing monopropellant compositions based on aqueous solutions of high-energy nitrogen-containing substances as the main propellant for low-thrust engines, for example, for meteorological rockets, for upper-stage engines, and in spacecraft control engine systems. This paper presents an approach that considers the selection and justification of ingredients based on renewable energy sources, the analysis being carried out primarily from standpoint of the availability of propellant components and their safety and energy efficiency. It is proposed that the energy of unitary reducing agent – oxidizer chemical propellants (energy-saturated compositions) be used as an alternative source. The development of nonhydrocarbon nitrogen-containing alternative energy sources with the possibility of their conversion and accumulation into the planetary nitrogen, oxygen, and water cycles is an urgent problem. The paper presents detailed information on propellant mixtures of nitrogen-containing substances as oxidizers and considers a number of reducing agents, such as alcohols, amides, etc. in composition with high-energy additives (aluminum, magnesium). The calculated results obtained meet the objectives and demonstrate that the compositions considered can be used as the main propellant for low-thrust engines. The advantages of the new propellant technology: availability, a low cost, produceability, environmental friendliness, a relatively low toxicity, and, primarily, a simpler design of the propulsion system and launch equipment. The proposed propellant composition, which is under test, is planned for use in the sustainer engines of ultralight suborbital rockets with the possibility of further development to an orbital rocket system.

APA, Harvard, Vancouver, ISO, and other styles

49

Zapata Usandivaras, José Felix, Annafederica Urbano, Michael Bauerheim, and Bénédicte Cuenot. "Data Driven Models for the Design of Rocket Injector Elements." Aerospace 9, no. 10 (October 12, 2022): 594. http://dx.doi.org/10.3390/aerospace9100594.

Full text

Abstract:

Improving the predictive capabilities of reduced-order models for the design of injector and chamber elements of rocket engines could greatly improve the quality of early rocket chamber designs. In the present work, we propose an innovative methodology that uses high-fidelity numerical simulations of turbulent reactive flows and artificial intelligence for the generation of surrogate models. The surrogate models that were generated and analyzed are deep learning networks trained on a dataset of 100 large eddy simulations of a single-shear coaxial injector chamber. The design of experiments was created considering three design parameters: chamber diameter, recess length, and oxidizer–fuel ratio. The paper presents the methodology developed for training and optimizing the data-driven models. Fully connected neural networks (FCNNs) and U-Nets were utilized as surrogate-modeling technology. Eventually, the surrogate models for the global quantity, average, and root mean square fields were used in order to analyze the impact of the length of the post’s recess on the performances obtained and the behavior of the flow.

APA, Harvard, Vancouver, ISO, and other styles

50

Abada, Omar, Abderahim Abada, and Ahmed Abdallah El-Hirtsi. "Effect of bipropellant combustion products on the rocket nozzle design." Mechanics & Industry 21, no. 5 (2020): 515. http://dx.doi.org/10.1051/meca/2020064.

Full text

Abstract:

The focus of this research work is to investigate numerically the effect of adding the gas on the design and performance of axisymmetric MLN nozzles. A FORTRAN code was developed to design this nozzle using the characteristics method (MOC) at high temperature. The thermochemical and combustion studies of the most used liquid propellants on the satellites and launch vehicles allow to known all gases. Four engines are investigated: Ariane 5 (Vulcain 2), Ariane-5 upper stage engine (Aestus), Zenit first stage (RD-170) and Falcon 9 upper stage (Raptor). Thermodynamic analysis of parameters design MLN (such as length, Mach number, mass, thrust coefficient) was conducted. The comparison shows that the presence of 50% of H2O gas in combustion species increases the nozzle design parameters (diatomic gas including air) in the order of 25%. On the other hand, the existence of CO2 gas considerably increases approximately 35% the length and the exhaust radius. These rise depend on gases percentage in the combustion. The truncation method is applied in the MLN nozzles to optimize the thrust/weight ratio.

APA, Harvard, Vancouver, ISO, and other styles

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

Create a spot-on reference in APA, MLA, Chicago, Harvard, and other styles