Journal articles: 'Liquid Propellant Rocket'

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Ma, Jiaju. "Analysis of the characteristics of rocket propellant." Theoretical and Natural Science 5, no. 1 (May 25, 2023): 490–95. http://dx.doi.org/10.54254/2753-8818/5/20230296.

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Rocket propellant is an important part of the rocket. Solid or liquid propellant will burn in the engine combustion chamber, and then a large amount of high-pressure gas will be generated. High-pressure gas will be ejected from the engine nozzle at a high speed, generating a reaction force on the rocket, so that the rocket will advance in the opposite direction of the gas injection. This paper mainly analyzes the advantages and disadvantages of current propellants and conceives a theoretically feasible propellant selection method. The main method of research is to calculate the theory of each propellant and make diagrams. Some of the research data were based on existing research reports. Each fuel has its own characteristics, they have one aspect of excellent ability, such as heat conduction, high thrust, high reliability. This paper summarizes the characteristics of current propellants and provides a more convenient query for future researches.

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Junqueira Pimont, Lia, Paula Cristina Gomes Fernandes, Luiz Fernando de Araujo Ferrão, Marcio Yuji Nagamachi, and Kamila Pereira Cardoso. "Study on the Mechanical Properties of Solid Composite Propellant Used as a Gas Generator." Journal of Aerospace Technology and Management, no. 1 (January 21, 2020): 7–10. http://dx.doi.org/10.5028/jatm.etmq.65.

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A gas generating propellants are used as initiators of liquid rocket propellants turbopumps and have as desired characteristic a high-volume production of low-temperature gas. In this context, some formulations of composite propellant containing polyurethane (based on liquid hydroxyl-terminated polybutadiene), guanidine nitrate, ammonium perchlorate, and additives were evaluated and characterized in order to verify their potential as gas generator propellant, as well as to evaluate the influence of additives on mechanical properties. The formulations were prepared, analyzed, and tested for mechanical properties.

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Abdelraouf, A. M., O. K. Mahmoud, and M. A. Al-Sanabawy. "Thrust termination of solid rocket motor." Journal of Physics: Conference Series 2299, no. 1 (July 1, 2022): 012018. http://dx.doi.org/10.1088/1742-6596/2299/1/012018.

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Abstract Rocket motors are engines that create the necessary thrust for the rocket motion. There are different types of rocket motors based on the propellant state, such as solid propellant rocket motors, liquid propellant rocket motors, and hybrid propellant rocket motors. One of the biggest disadvantages of solid propellant rocket motors, in comparison to liquid and hybrid propellant rocket motors, is that they are extremely difficult to extinguish, necessitating the use of specific devices. This paper reviews various ways for thrust termination such as fluid injection, rapid increase in throat area, and sudden opening of an additional port at the forward section of the motor, which increases the depressurization rate (dp/dt) required for extinguishing. The rate of depressurization varies depending on propellant components, combustion pressure, and exhaust pressure, and may be investigated using experimental approaches. The change in the critical area for a motor can be predicted by using MATLAB code to ensure the complete extinguishing by decreasing the pressure under the deflagrationlimit with high depressurization rate.

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Czerwińska, Magdalena, and Piotr Prasuła. "STUDY OF THERMO-MECHANICAL PROPERTIES OF AGED HOMOGENEOUS SOLID ROCKET PROPELLANT ACCORDING TO STANAG REQUIREMENTS." PROBLEMY TECHNIKI UZBROJENIA 145, no. 1 (May 15, 2018): 47–63. http://dx.doi.org/10.5604/01.3001.0012.1325.

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Proper operation of rocket motor depends significantly on thermo-mechanical properties of propellant used. For this reason it is important that characteristics of a particular propellant versus the time and temperature pass a thorough investigation to assess its operation at different conditions. The paper illustrates investigations of ageing process influencing thermo-mechanical properties of homogeneous rocket propellant. A selected type of rocket propellant was subjected to accelerated ageing in conditions specified in AOP-48 document to establish in the next step its thermal and mechanical characteristics (between all the temperature of glass transition and decomposition). The ageing of propelling explosives causes the reduction of stabiliser content deciding about thermo-mechanical properties of propellant and for that the percentage of effective stabiliser and its loss were identified by liquid chromatography HPLC. Thermal properties were investigated by differential scanning calorimetry. Thermal analyses were carried out according to STANAG 4515. Mechanical characteristics were tested by dynamic mechanical analysis (DMA) in line with STANAG 4540.

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ZAGANESCU, Nicolae-Florin, Rodica ZAGANESCU, and Constantin-Marcian GHEORGHE. "Wernher Von Braun’s Pioneering Work in Modelling and Testing Liquid-Propellant Rockets." INCAS BULLETIN 14, no. 2 (June 10, 2022): 153–61. http://dx.doi.org/10.13111/2066-8201.2022.14.2.13.

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This paper presents a view on how Dr. Wernher Von Braun laid the basis for realistic modelling and testing liquid-propellants rockets, by his PhD Thesis – a secret document in 1934, which remained classified until 1960. Understanding that better mathematical modelling is needed if these rockets are to become spaceflight vehicles, he clarified in his thesis essential issues like: maximum achievable rocket speed; Laval nozzle thrust gain; polytropic processes in the combustion chamber and nozzle; influence of equilibrium and dissociation reactions; original measurement systems for rockets test stand; engineering solutions adequate for series production of the combustion chamber – reactive nozzle assembly. The thesis provided a theoretical and experimental basis for a new concept of the rocket, having a lightweight structure; low tanks pressure; high-pressure pumps and injectors; low start speed; rocket stabilization by gyroscopic means or by active jet controls; longer engine burning time; higher jet speed. Numerous tests made even with a fully assembled rocket (the “Aggregate-I”), improved mathematical model accuracy (e.g., the maximum achievable altitude predicted for the “Aggregate-II” rocket was confirmed later in-flight tests).

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Yuan, Wen-Li, Lei Zhang, Guo-Hong Tao, Shuang-Long Wang, You Wang, Qiu-Hong Zhu, Guo-Hao Zhang, et al. "Designing high-performance hypergolic propellants based on materials genome." Science Advances 6, no. 49 (December 2020): eabb1899. http://dx.doi.org/10.1126/sciadv.abb1899.

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A new generation of rocket propellants for deep space exploration, ionic liquid propellants, with long endurance and high stability, is attracting more and more attention. However, a major defect of ionic liquid propellants that restricts their application is the inadequate hypergolic reactivity between the fuel and the oxidant, and this defect results in local burnout and accidental explosions during the launch process. We propose a visualization model to show the features of structure, density, thermal stability, and hypergolic activity for estimating propellant performances and their application abilities. This propellant materials genome and visualization model greatly improves the efficiency and quality of developing high-performance propellants, which benefits the discovery of new advanced functional molecules in the field of energetic materials.

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7

Lawrence, Lovell. "A LIQUID-PROPELLANT ROCKET MOTOR.*." Journal of the American Society for Naval Engineers 58, no. 4 (March 18, 2009): 642–45. http://dx.doi.org/10.1111/j.1559-3584.1946.tb02717.x.

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8

Cheng, Yuqiang, and Jianjun Wu. "Particle swarm algorithm-based damage-mitigating control law analysis and synthesis for liquid-propellant rocket engine." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 233, no. 10 (October 31, 2018): 3810–18. http://dx.doi.org/10.1177/0954410018806080.

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The damage-mitigating control is a novel technique to ameliorate the reliability and safety of liquid-propellant rocket engines by achieving an optimized trade-off level between overall dynamic performance of the liquid-propellant rocket engine and structural durability of some selected critical damageable components under the condition of no impact on the achievement of the launch and flight mission. Thus, it is needed to be solved for the damage-mitigating control that the global optimization of the best trade-off between the damage of the critical damageable components and the performance of rocket engine. The major challenge should focus on: (i) to construct model of a certain rocket engine system dynamics, critical components structural dynamics, and damage dynamics; (ii) to optimize open loop feed-forward control law based on liquid-propellant rocket engine system dynamic model, structural and damage dynamics model, by using particle swarm optimization algorithm; (iii) to synthesize an intelligent damage-mitigating control system using the optimized open loop control law. In this paper, synthesis procedure of damage mitigation is introduced; structure and damage dynamic model of damageable components are formulated. The results of the simulation computation show that the synthesized control laws are implemented and achieve the effect of damage mitigating for the liquid-propellant rocket engine. It can provide important theoretical and practical value not only for improving the safety and reliability of the liquid-propellant rocket engine, but also for the complex thermo-flow-mechanical systems such as airplane engines, automobile engines, and fossil-fueled power plant because their service life is very critical too.

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Palacz, Tomasz, and Jacek Cieślik. "Experimental Study on the Mass Flow Rate of the Self-Pressurizing Propellants in the Rocket Injector." Aerospace 8, no. 11 (October 26, 2021): 317. http://dx.doi.org/10.3390/aerospace8110317.

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High vapor pressure propellants such as nitrous oxide are widely used in experimental hybrid and liquid rockets as they can be used in a self-pressurization mode, eliminating the need for external pressurization or pumps and simplifying the design of the rocket system. This approach causes the two-phase flow in the feed system and the injector orifices, which cannot be easily modeled and accounted for in the design. A dedicated test stand has been developed to better understand how the two-phase flow of the self-pressurizing propellant impacts the mass flow characteristics, enabling the simulation of the operating conditions in the rocket engine. The injectors have been studied in the range of ΔP. The flow regimes have been identified, which can be predicted by the SPI and HEM models. It has been shown that the two-phase flow quality upstream of the injector may impact the discharge coefficient in the SPI region and the accuracy of the HEM model. It has been found that the transition to the critical flow region depends on the L/D ratio of the injector orifice. A series of conclusions can be drawn from this work to design the rocket injector with a self-pressurizing propellant to better predict the mass flow rate and ensure stable combustion.

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Khan, Tajwali, and Ihtzaz Qamar. "Factors Affecting Characteristic Length of the Combustion Chamber of Liquid Propellant Rocket Engines." July 2019 38, no. 3 (July 1, 2019): 729–44. http://dx.doi.org/10.22581/muet1982.1903.16.

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Optimum characteristic length of the combustion chamber of liquid rocket engine is very important to get higher energy from the liquid propellants. Characteristic length is defined by the time required for complete burning of fuel. Combustion reactions are very fast and combustion is evaporation dependent. This paper proposes fuel droplet evaporation model for liquid propellant rocket engine and discusses the factors which can affect the required size of characteristic length of the combustion chamber based on proposed model. The analysis is performed for low temperature combustion chamber. A computer code based on proposed model is generated, which solve analytical equations to calculate combustion chamber characteristic length under various input conditions. The analysis shows that characteristic length is affected by combustion chamber temperature, pressure, fuel droplet diameter, chamber diameter, mass flow rate of propellants and relative velocity of the droplet in the combustion chamber.

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11

Popov, Pavel P., William A. Sirignano, and Athanasios Sideris. "Propellant Injector Influence on Liquid-Propellant Rocket Engine Instability." Journal of Propulsion and Power 31, no. 1 (January 2015): 320–31. http://dx.doi.org/10.2514/1.b35400.

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12

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.

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13

Bazarov, Vladimir G., and Vigor Yang. "Liquid-Propellant Rocket Engine Injector Dynamics." Journal of Propulsion and Power 14, no. 5 (September 1998): 797–806. http://dx.doi.org/10.2514/2.5343.

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14

Лазарева, Ю. И., С. В. Клименко, А. В. Кулик, and И. В. Лазарев. "АНАЛИЗ СОВРЕМЕННОГО СОСТОЯНИЯ И ПЕРСПЕКТИВЫ РАЗВИТИЯ РАКЕТНЫХ ДВИГАТЕЛЕЙ ДЛЯ ИССЛЕДОВАНИЯ ДАЛЬНЕГО КОСМОСА." System design and analysis of aerospace technique characteristics 27, no. 2 (May 17, 2022): 50–58. http://dx.doi.org/10.15421/471923.

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The article analyzes the current state and development prospects of rocket engines for space exploration. Currently, for flights to other planets, not to mention the stars, the use of liquid-propellant and solid-propellant rocket engines is becoming increasingly unprofitable, although many rocket engines have been developed. Thus, to reach manned planes even the nearest planets, it is necessary to develop rocket launchers on engines operating on principles different from chemical propulsion systems. The most promising in this regard are electric, laser and nuclear rocket engines, as well as hybrid rocket engines such as solid-state nuclear rocket engine (NRE) and electric jet engine (ERE) or gas-phase NRE and ERE.

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15

Nie, Yao, Yuqiang Cheng, and Jianjun Wu. "Liquid-propellant rocket engine online health condition monitoring base on multi-algorithm parallel integrated decision-making." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 231, no. 9 (July 12, 2016): 1621–33. http://dx.doi.org/10.1177/0954410016656878.

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This study reports multi-algorithm parallel integrated decision-making for liquid-propellant rocket engine online health condition monitoring to improve reliability and safety, especially for next-generation reusable engines. Fusing multi-algorithm detection information to judge liquid-propellant rocket engine condition is multi-algorithm parallel integrated decision-making main task, and multi-algorithm judgment problem is its central issue; i.e. how to make a global judgment from judgment results of different fault detection methods. Considering opportune fault detection, adequate rocket engine information exploitation, and reliable condition judging, the multi-algorithm parallel integrated decision-making framework for problem definition is presented along with a multi-algorithm parallel integrated decision-making judgment model. For more reliable, efficient global judgment, a method based on the Bayes’ risk function integrating multi-algorithm prior information is adopted. The proposed approach is validated with liquid-propellant rocket engine ground testing data. The results show that the multi-algorithm parallel integrated decision-making judgment model gives very effective and reliable performance relative to the voting method, successfully solving multi-algorithm judgment problems and meeting practical engineering needs.

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Orlin, Sergei A. "Use of cryogenic components of propellants for liquid-propellant rocket engines and in life support systems of manned space vehicles." MATEC Web of Conferences 324 (2020): 01005. http://dx.doi.org/10.1051/matecconf/202032401005.

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The cited materials show the use of oxygen, hydrogen, liquefied natural gases (methane) and fluorine as components of the fuel for liquid-propellant rocket engines (LRE). The reasons for the need to use oxygen as an oxidizing agent are indicated. The advantages and disadvantages are disclosed from the point of view of using the listed components as fuel elements for liquid-propellant rocket engines. The issues of ecology when using the considered fuels are reviewed. Shown not only the use of cryogenic components as fuel for LRE, but also in life support systems in manned spacecraft in space research.

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17

Zhang, Yulin, Jianjun Wu, Minchao Huang, Hengwei Zhu, and Qizhi Chen. "Liquid-Propellant Rocket Engine Health-Monitoring Techniques." Journal of Propulsion and Power 14, no. 5 (September 1998): 657–63. http://dx.doi.org/10.2514/2.5327.

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18

Lux, Johannes, Dmitry Suslov, and Oskar Haidn. "On porous liquid propellant rocket engine injectors." Aerospace Science and Technology 12, no. 6 (September 2008): 469–77. http://dx.doi.org/10.1016/j.ast.2007.11.004.

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Yang, Wei, and Bing Sun. "Numerical Simulation of Liquid Film in a Liquid Oxygen/Rocket Propellant 1 Liquid Rocket." Journal of Thermophysics and Heat Transfer 26, no. 2 (April 2012): 328–36. http://dx.doi.org/10.2514/1.t3759.

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20

Sidlerov, D. A., and S. A. Fedorov. "Numerical Investigation of Work Cycle Characteristics in the Combustion Chamber of a Lox/Methane Liquid-Propellant Rocket Engine Featuring Reductant Power Gas Combustion." Herald of the Bauman Moscow State Technical University. Series Mechanical Engineering, no. 2 (141) (June 2022): 43–53. http://dx.doi.org/10.18698/0236-3941-2022-2-43-53.

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We performed a numerical investigation of cumulative efficiency and the structure in detail concerning the working process in the combustion chamber of a lox/methane liquid-propellant rocket engine operating in steady-state, boosted and throttled modes. In order to do it, we used tools developed by JSC SSC "Center Keldysh", that is, physical and mathematical models, numerical methods and software packages for numerical simulation of two-phase turbulent flows with combustion in liquid-propellant engine combustion chambers. The paper presents numerical simulation and investigation results concerning the specifics of fuel component flows, their mixing and combustion in the combustion chamber of a lox/methane liquid-propellant rocket engine using staged combustion cycle with reductant gas in steady-state, boosted (117 % by thrust) and throttled (30 % by thrust) operation modes. We performed a comparative analysis of work cycle parameters in combustion chambers at different fuel component consumption rates and pressure levels. The paper shows that the boosted mode increases the interaction of fuel jets, which intensifies mixing and burnout processes, while the deep throttling mode decreases the mixing and fuel burnout amplitudes as compared to the steady-state mode. The numerical simulation results may be used to investigate fuel combustion processes in combustion chambers of promising liquid-propellant rocket engines at the stages of development, design and refinement

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Leto, Angelo. "Investigation of a Radial Turbines Compatibility for Small Rocket Engine." E3S Web of Conferences 197 (2020): 11009. http://dx.doi.org/10.1051/e3sconf/202019711009.

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In the radial turbine preliminary design for an expander rocket engine, a comparison was made with axial turbine used in Pratt & Whitney RL10 engine. One of the primary requirements of a liquid propellant rocket engine is the generation of a high thrust, which depends on both the mass flow rate of the propellant and the pressure in the thrust chamber. In expander-cycle engines, which are the subject of the present study, the liquid propellant is first compressed using centrifugal turbo-pumps, then it is used to cool the combustion chamber and the nozzle and, once vaporized, it flows through the turbines used to drive the turbo-pumps. The aim was to demonstrate the greater efficiency of the radial turbine with a reduction of the pressure ratio with respect to the axial turbine.

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LEE, Yang-Suk, and Jun Hwan JANG. "The design and performance on 200N-class bipropellant rocket engine using decomposed H2O2 and Kerosene." INCAS BULLETIN 11, no. 3 (September 9, 2019): 99–110. http://dx.doi.org/10.13111/2066-8201.2019.11.3.9.

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Mono-propellant thrusters are widely utilized in satellites and space launchers. In many cases, they are using hydrazine as a propellant. However, hydrazine has high toxicity and high risks in using for launch campaign. Recently, low-toxic (green) propellant is being highlighted as a replacement for hydrazine. In this paper, 200N bi-propellant engine using hydrogen peroxide/kerosene was designed/manufactured, and the spray or atomization characteristic and inflation pressure were determined by cold flow test, and combustion and pulse tests in a single cycle same as previous methods were conducted. As uniformly supplying hydrogen peroxide through plate-type orifice to a catalyst bed, the hot gas was created as a reaction with hydrogen and catalyst. And then, it was confirmed that the ignition is possible on the wide range of O/F ratio without additional ignition source. The liquid rocket engine with bi-propellant of hydrogen peroxide/kerosene and design/test methods which developed in this study are expected to be utilized as an essential database for designing of the ignitor/injector of bi-propellant liquid rocket engine using hydrogen peroxide/kerosene with high-thrust/performance in near future.

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23

Strelnikov, G. A., A. D. Yhnatev, N. S. Pryadko, and S. S. Vasyliv. "Gas flow control in rocket engines." Technical mechanics 2021, no. 2 (June 29, 2021): 60–77. http://dx.doi.org/10.15407/itm2021.02.060.

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In the new conditions of application of launch vehicle boosters, space tugs, etc., modern rocket engines often do not satisfy the current stringent requirements. This calls for fundamental research into processes in rocket engines for improving their efficiency. In this regard, for the past 5 years, the Department of Thermogas Dynamics of Power Plants of the Institute of Technical Mechanics of the National Academy of Sciences of Ukraine and the State Space Agency of Ukraine has conducted research on gas flow control in rocket engines to improve their efficiency and functionality. Mechanisms of flow perturbation in the nozzle of a rocket engine by liquid injection and a solid obstacle were investigated. A mathematical model of supersonic flow perturbation by local liquid injection was refined, and new solutions for increasing the energy release rate of the liquid were developed. A numerical simulation of a gas flow perturbed by a solid obstacle in the nozzle of a rocket engine made it possible to verify the known (mostly experimental) results and to reveal new perturbation features. In particular, a significant increase in the efficiency of flow perturbation by an obstacle in the transonic region was shown up, and some dependences involving the distribution of the perturbed pressure on the nozzle wall, which had been considered universal, were refined. The possibility of increasing the efficiency of use of the generator gas picked downstream of the turbine of a liquid-propellant rocket engine was investigated, and the advantages of a new scheme of gas injection into the supersonic part of the nozzle, which provides both nozzle wall cooling by the generator gas and the production of lateral control forces, were substantiated. A new concept of rocket engine thrust vector control was developed: a combination of a mechanical and a gas-dynamic system. It was shown that such a thrust vector control system allows one to increase the efficiency and reliability of the space rocket stage flight control system. A new liquid-propellant rocket engine scheme was developed to control both the thrust amount and the thrust vector direction in all planes of rocket stage flight stabilization. New approaches to the process organization in auxiliary elements of rocket engines on the basis of detonation propellant combustion were developed to increase the rocket engine performance.

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24

Huh, Jeongmoo, Botchu V. S. Jyoti, Yongtae Yun, M. N. Shoaib, and Sejin Kwon. "Preliminary Assessment of Hydrogen Peroxide Gel as an Oxidizer in a Catalyst Ignited Hybrid Thruster." International Journal of Aerospace Engineering 2018 (December 30, 2018): 1–14. http://dx.doi.org/10.1155/2018/5630587.

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In regard to propulsion system applications, the stability of liquid propellants in long-term storage is of increasing importance, and this had led to a greater interest in gelation technology. As part of a preliminary test to determine the feasibility of using a gel propellant in a rocket with a catalyst bed, a hybrid rocket with a catalyst reactor using a gel propellant as an oxidizer was tested for the first time in this study. Experiments were conducted with two different oxidizers: one with liquid phase hydrogen peroxide and the other with gel phase hydrogen peroxide, as well as high-density polyethylene as fuel for a 250 N class hybrid thruster performance test. The thruster was designed with the catalyst ignition system, and a catalyst was manufactured to be inserted into the catalyst reactor to facilitate oxidizer decomposition. While the test result with neat hydrogen peroxide indicated sufficient decomposition efficiency using a manganese dioxide/alumina catalyst and successful autoignition of the fuel via the decomposed product, gel hydrogen peroxide exhibited insufficient decomposition and there were difficulties in operating the thruster as a part of the catalyst was covered in the gelling agent. This preliminary study identifies the potential challenges of using a gel phase oxidizer in a catalyst ignited hybrid thruster and discusses the technical issues that should be addressed in regard to a gel propellant hybrid thruster design with a catalyst reactor.

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Nie, Yao, Yuqiang Cheng, and Jianjun Wu. "Dynamic cloud back-propagation networks and its application in fault diagnostic for liquid-propellant rocket engines." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 232, no. 3 (December 16, 2016): 583–94. http://dx.doi.org/10.1177/0954410016683413.

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Fault diagnosis for liquid-propellant rocket engines often faces a lack of prior knowledge or insufficient sampling data, and thus becomes a decision-making problem with uncertain information sources. In this paper, a method based on dynamic cloud back-propagation networks is proposed. This uses cloud theory to synthetically combine randomness and fuzziness. In this work, a cloud model and back-propagation neural network are synthetically combined in series. A cloud transformation is used to identify the network structure and extract the features of the cloud model. Simultaneously, a unit-delay step is introduced into the input layer to describe the dynamic behaviour during the engine working process. The proposed fault diagnosis method for liquid-propellant rocket engines is verified using the actual data. The results confirm that the proposed method accurately recognizes all three relevant failure modes. Further, randomness associated with the measurement process and ambient noises are simulated by adding random noise to the test conditions. Simulation results demonstrate that the method correctly detects and classifies faults according to the principles of sustainability, indicating a high robustness towards noise. The proposed method has a single-step operating time of 1.24 × 10−4 s, satisfying the real-time requirements for fault diagnosis in liquid-propellant rocket engines.

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Casiano, Matthew J., James R. Hulka, and Vigor Yang. "Liquid-Propellant Rocket Engine Throttling: A Comprehensive Review." Journal of Propulsion and Power 26, no. 5 (September 2010): 897–923. http://dx.doi.org/10.2514/1.49791.

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27

TSUTSUMI, Seiji, Takanori HAGA, and Taro SHIMIZU. "Combustion Simulation in Liquid-propellant Rocket Engine Development." Journal of the Visualization Society of Japan 41, no. 162 (2021): 19–20. http://dx.doi.org/10.3154/jvs.41.162_19.

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Ha, Seong-Up, Oh-Sung Kwon, Jung-Ho Lee, Byoung-Hun Kim, Sun-Il Kang, Sang-Yeop Han, In-Hyun Cho, and Dae-Sung Lee. "Determination of Cyclogram for Liquid-Propellant Rocket Engine." International Journal of Aeronautical and Space Sciences 3, no. 2 (November 30, 2002): 59–66. http://dx.doi.org/10.5139/ijass.2002.3.2.059.

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29

Salvador, Nicolas M. C., Marcelo M. Morales, Carlos E. S. S. Migueis, and Demétrio Bastos-Netto. "Numerical simulation of a liquid propellant rocket motor." Journal of Thermal Science 10, no. 1 (March 2001): 83–86. http://dx.doi.org/10.1007/s11630-001-0015-8.

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Wu, Jianjun. "Liquid-propellant rocket engines health-monitoring—a survey." Acta Astronautica 56, no. 3 (February 2005): 347–56. http://dx.doi.org/10.1016/j.actaastro.2004.05.070.

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Vasques, Bruno Barreto, and Luiz Carlos Gadelha de Souza. "Investigation of the Lox/Alcohol Propulsion System." Applied Mechanics and Materials 319 (May 2013): 427–32. http://dx.doi.org/10.4028/www.scientific.net/amm.319.427.

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This work describes the design of a liquid propellant rocket engine using non-toxic, low-cost propellants, for a second generation space transportation system. The primary goals of this effort were to identify the most attractive fuel and system design approach for sizing a workhorse engine and to determine technology advancements that are needed to provide subsequent system development. The emphasis of the analysis was directed toward propellant perfomance and engine characterization. On the basis of operational and cost criteria, ethanol was determined to be the best fuel candidate. Because of the high heat flux conditions at high chamber pressures, regenerative cooling was chosen as the preferred method of wall protection, leaving film cooling as a supplementary chamber cooling method. A finite element analysis was carried out to evaluate thrust chamber reliability on steady working operation and hydraulic test mode. The study contributed to establish a better understanding of propellant potentialities and to investigate the best means of utilizing these capabilities in typical engine designs.

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Pylypenko, O. V., O. O. Prokopchuk, S. I. Dolgopolov, O. D. Nikolayev, N. V. Khoriak, V. Yu Pysarenko, I. D. Bashliy, and S. V. Polskykh. "Mathematical modelling of start-up transients at clustered propulsion system with POGO-suppressors for CYCLON-4M launch vehicle." Kosmìčna nauka ì tehnologìâ 27, no. 6 (2021): 3–15. http://dx.doi.org/10.15407/knit2021.06.003.

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Liquid-propellant rocket propulsion systems of the first stages of launch vehicles of medium, heavy, and super-heavy class usually include POGO-suppressors, which are one of the most widely used methods to eliminate launch vehicle longitudinal structural vibrations (POGO phenomena). However, until now, the theoretical studies and analysis of the effect of the POGO-suppressors’ installation in the feedlines of main liquid rocket engines on transient processes in systems during rocket engine starting have not been carried out due to the complexity of such analysis and the lack, first of all, reliable nonlinear models of cavitation phenomena in rocket engine pumps. A mathematical model for the start-up of a clustered rocket propulsion of the Cyclone-4M launch vehicle has been developed that takes into account the low-frequency dynamics of the POGO-suppressors and the asynchronous start-up timeline sequences of the rocket engines. The first stage of the launch vehicle propulsion system includes four RD-870 rocket engines. A nonlinear mathematical model of low-frequency dynamic processes of the POGO-suppressor with bellows separation of liquid and gaseous media is presented. A significant effect of cavitation in the pumps of engines and the POGO-suppressor installation to the LOX feedline on the propulsion system dynamic gains is shown. Based on the developed mathematical model of the clustered rocket propulsion start-up, the studies of the Cyclone-4M main engines’ start-up transients were carried out. The asynchronous start-up timeline sequences of the rocket engine and the places of installation of the POGO-suppressors in the LOX feedline branches to the RD-870 rocket engine – near the general feedline collector as standard placement or directly at the entrance to the engines – were investigated. The analysis of start-up transients in the oxidizer feed system of the considered propulsion (the time dependences of the flowrate and pressure at the engine inlet) showed the following. Firstly, while the synchronous start-up of the engines, the installation of the POGO-suppressors near the feedline collector makes it possible to eliminate all engine inlet overpressures that exist in the rocket propulsion system in case of the absence of the POGO-suppressors. Secondly, the RD-870 engine asynchronous start-up operation affects negatively the time dependences of the propellant flowrate and pressure at the engine inlet if the POGO-suppressors are located near the feedline collector. So, in the propulsion system’s start-up timeline interval 0.95 s - 1.35 s, for some computational variants of the initial moments of the engine operation start, an abnormally large drop in the LOX flow rate and the overpressures at the engine inlet is observed. The asynchronous start-up of the RD-870 engines with the installation of the POGO-suppressors at the engine inlet does not significantly change the start-up transients compared to the synchronous starting of the engines. Thirdly, thus, it is shown that the installation of the POGO-suppressors both at the engine inlet and at the RD-870 branches near the collector has a significant positive effect on the quality of start-up transient processes for the main engines of the 1st stage of the Cyclone-4M launch vehicle. Placing the POGO-suppressors at the engine inlets is not standard and is considered without reference to the propulsion system layout. Nevertheless, the POGO-suppressors installed at the inlet to the engines are an effective means of preventing overshoots and dips in the parameters of the liquid-propellant rocket engine, including the conditions of asynchronous starting of the liquid rocket engines in the clustered propulsion system. The results obtained can be used in mathematical modeling of the start-up of the first stage propulsion system either for multistage sustainer rockets used in parallel with booster rockets or for the clustered multi-engine rocket propulsion system containing POGO-suppressors.

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Kozin, V. S. "Effect of the thermal and gas-dynamic properties of solid rocket propellant particles on the propellant combustion rate." Technical mechanics 2021, no. 1 (April 30, 2021): 63–67. http://dx.doi.org/10.15407/itm2021.01.063.

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The aim of this work is to eliminate the explosion possibility of a rocket engine that operates on a fast-burning solid propellant. The problem is considered by analogy with experiments conducted earlier. Various ways to increase the propellant combustion rate are presented. Examples of how the solid propellant combustion rate depends on the metal fuel and the oxidizer particle size are given. It is shown that unstable combustion of a solid propellant at high combustion chamber pressures is due to unstable combustion of the gas phase in the vicinity of the bifurcation point. Zeldovich’s theory of nonstationary powder combustion is applied to analyzing the explosion dynamics of the Hrim-2 missile’s solid-propellant sustainer engine. This method of analysis has not been used before. The suggested version that this phenomenon is related to the aluminum particle size allows one to increase the combustion rate in the combustion chamber of a liquid-propellant engine, thus avoiding the vicinity of the bifurcation point. The combustion of solid propellants differing in aluminum particle size is considered. The metal fuel and the oxidizer particle sizes most optimal in terms of explosion elimination are determined and substantiated. The use of submicron aluminum enhances the evaporation of ammonium perchlorate due to the infrared radiation of aluminum particles heated to an appropriate radiation temperature. This increases the gas inflow into the charge channel, thus impeding the suppression of ammonium perchlorate sublimation by a high pressure, which is important in the case where the engine body materials cannot withstand a high pressure in the charge channel. This increases the stability and rate of solid propellant combustion. It is shown that the Hrim-2 missile’s solid propellant cannot be used in the Hran missile. The combustion rate is suggested to be increased by using fine-dispersed aluminum in the solid propellant.

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Wang, Guo Qing, Jun Bo Jia, and Xu Yuan Li. "Research on Feature Selection Based on Improved Particle Swarm Optimization." Advanced Materials Research 591-593 (November 2012): 2651–54. http://dx.doi.org/10.4028/www.scientific.net/amr.591-593.2651.

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Feature selection is one of key technologies for fault diagnosis. Especially for high dimensional data, Feature selection can not only find the feature subset with sufficient information, but also improve the classification accuracy and efficiency. In order to decrease the number of diagnosis parameter in fault diagnosis of Liquid-propellant Rocket Engine, the paper proposes one feature selection method based on improved particle swarm optimization, the method applies the quantum evolution thoughts to PSO. The particle is restricted in the range from -π/2 to 0, so the particle can correspond to the quantum angle. The parameter optimization function is designed. The improved algorithm can decrease the number of parameter in fault diagnosis of Liquid-propellant Rocket Engine from 25 to 6.

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Bandyopadhyay, Atri, and Ankit Kumar Mishra. "Comparative Study on Hybrid Rocket Fuels for Space Launch Vehicles Moving in Higher Orbits." 4 1, no. 4 (December 2, 2022): 13–19. http://dx.doi.org/10.46632/jame/1/4/3.

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The current research is focused on understanding the propulsive parameters of hybrid rocket motors. A comparative study is prepared from various research papers. The fuels paraffin wax, hydroxyl-terminated polybutadiene (HTPB) and polymethyl methacrylate (PMMA) with added additives (Al/Mg) were combined with two oxidizers, liquid oxygen (LOX), nitrous oxide (N2O). The propulsive parameters examined were the combustion efficiency, combustion or adiabatic flame temperature, characteristics velocity and regression rate. The propellant pair paraffin-N2O provided the highest performance for all parameters studied. This study provides an advantageous propellant option for future rocket propulsion based on a comparative investigation

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Nikolayev, O. D., I. D. Bashliy, N. V. Khoriak, and S. I. Dolgopolov. "Evaluation of the high-frequency oscillation parameters of a liquid-propellant rocket engine with an annular combustion chamber." Technical mechanics 2021, no. 1 (April 30, 2021): 16–28. http://dx.doi.org/10.15407/itm2021.01.016.

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The high-frequency instability (HF instability) of a liquid-propellant rocket engine (LPRE) during static firing tests is often accompanied by a significant increase in dynamic loads on the combustion chamber structure, often leading to the chamber destruction. This dynamic phenomenon can also be extremely dangerous for the dynamic strength of a liquid-propellant rocket engine with an annular combustion chamber. Computation of the parameters of acoustic combustion product oscillations is important in the design and static firing tests of such rocket engines. The main aim of this paper is to develop a numerical approach to determining the parameters of acoustic oscillations of combustion products in annular combustion chambers of liquid-propellant rocket engines taking into account the features of the configuration of the combustion space and the variability of the physical properties of the gaseous medium depending on the axial length of the chamber. A numerical approach is proposed. The approach is based on mathematical modeling of natural oscillations of a “shell structure of an annular chamber – gas” coupled dynamic system by using the finite element method. Based on the developed finite-element model of coupled spatial vibrations of the structure of the annular combustion chamber and the combustion product oscillations, the oscillation parameters of the system under consideration (frequencies, modes, and effective masses) for its dominant acoustic modes, the vibration amplitudes of the combustion chamber casing, and the amplitudes of its vibration accelerations can be determined. The operating parameters of the liquid-propellant rocket engine potentially dangerous for the development of thermoacoustic instability of the working process in the annular combustion chamber can be identified. For the numerical computation of the dynamic gains (in pressure) of the combustion chamber, a source of harmonic pressure excitation is introduced to the finite element model of the dynamic system “shell structure of an annular configuration – gas” (to the elements at the start of the chamber fire space). The developed approach testing and further analysis of the results were carried out for an engine with an annular combustion chamber (with a ratio of the outer and inner diameters of 1.5) using liquid oxygen – methane as a propellant pair. The system shapes and frequencies of longitudinal, tangential and radial modes are determined. It is shown that the frequency of the first acoustic mode in the case of a relatively low stiffness of the combustion chamber casing walls can be reduced by 40 percent in comparison with the frequency determined for a casing with rigid walls.

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Khan, Sohaib, Muhammad Umer Sohail, Ihtzaz Qamar, Muzna Tariq, and Raees Fida Swati. "Effect of Secondary Combustion on Thrust Regulation of Gas Generator Cycle Rocket Engine." Applied Sciences 12, no. 20 (October 19, 2022): 10563. http://dx.doi.org/10.3390/app122010563.

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Thrust regulation is applied to maintain the performance of the liquid propellant rocket engine. The thrust level of a rocket engine can be readily controlled by adjusting the number of propellants introduced into the combustion chamber. In this study, a gas generator design is proposed in which thrust regulation is maintained by performing secondary combustion in the divergent section of the nozzle of a gas generator. Tangential and normal injection techniques have also been studied for better combustion analyses. A normal injection technique is used for the experiment and CFD results are validated with the experimental data. Chemical equilibrium analyses are also performed by minimizing Gibbs free energy with the steepest descent method augmented by the Nelder–Mead algorithm. These equilibrium calculations give the combustion species as obtained through the CFD results. Performance evaluation of the rocket engine, with and without secondary combustion in the gas generator, led to an increase of 42% thrust and 46.15% of specific impulse with secondary combustion in the gas generator.

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Feng, Song Jiang, Hao Bo He, Xue Liu, Bo He, and Wan Sheng Nie. "Investigation of the Evaporation Processes of Gel Propellant Droplets." Advanced Materials Research 146-147 (October 2010): 753–56. http://dx.doi.org/10.4028/www.scientific.net/amr.146-147.753.

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Due to their high-performance and improved safety, gel propellants can be used in various boost motors and large launchers. The evaporation and combustion characteristics of gel propellants are the foundation for the gel-engine design. Especially, it is basal and important to study the gel droplet evaporation process. In this paper, the gel droplet evaporation model is developed to simulate the gel droplet evaporation process at first. Then the experiments to record the gel droplet evaporation process are conducted. During the droplet evaporation process, the decreased velocity of the droplet diameter increases gradually, whereas that of the droplet mass decreases gradually. The mass of both the liquid fuel and the gellant decreases gradually, however, the gellant mass concentration increases gradually and at the evaporation later stage the gellant mass is larger than that of the liquid fuel. The typical evaporation process characteristics captured by experiments are in reasonable agreement with the gel droplets evaporation mechanism. Especially, the “micro-burst” phenomenon of the gelled propellant may appears in rocket engines.

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Bortoloto, Evandro Rostirolla, Francisco Carlos Parquet Bizarria, and José Walter Parquet Bizarria. "Petri Nets Applied in Purge Algorithm Analysis for a Rocket Engine Test with Liquid Propellant." Aerospace 10, no. 3 (February 24, 2023): 212. http://dx.doi.org/10.3390/aerospace10030212.

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During the development stage of a space vehicle, instrumented tests are carried out on the ground to prove the operational capacity of each liquid-propellant rocket engine, which is installed in this type of vehicle. The task of elaborating a Test Bench project for a propulsion unit with this application is complex and involves several steps, one of these steps being related to the analysis of this bench capacity to meet the algorithms for the liquid-propellant rocket-engine full run of tests, which is considered fundamental for this project’s operational success. Due to the high costs involved in this project’s elaboration and execution, it is strategic to use computational resources to evaluate, by simulation, the main operational functionalities that are previously established for this bench to perform. In this context, this work presents a model proposal through Petri Nets to evaluate, by computer simulation, an architecture capacity that was designed for the Test Bench to meet an algorithm dedicated to the liquid-propellant pipelines purge during the run of hot tests with the liquid-propellant rocket engine. The method used in this work to carry out the simulation shows the operational response of each module of this architecture, in accordance with the steps contained in the purge algorithm, which allows for analyzing, for each event of the process, the Petri Nets properties, mainly those related to the conservativeness, liveliness, deadlock-type, and confusion-type conflicts. The simulation carried out with the proposed model allows for the portrayal of the physical architecture and the operational states of the purge system according to the steps foreseen in the algorithm, showing that the conservation property is met because the number of marks remains constant, the vivacity property is also met since all positions have been reached, and there is no mortal-type conflict, as the simulation is not stopped; only confusion-type conflict is identified, which was solved with the strategic insertion of resources in the model in order to fix crashes related to the competition for tokens in the transition-enabled entries. The satisfactory results obtained in these simulations suggest that the modules provided for this architecture are sufficient and appropriate for carrying out all the steps contained in the purge algorithm, which will minimize or even eliminate the disorders that may be caused by the presence of foreign elements in the propellant supply lines during the tests with the rocket engine.

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40

IN, Sehwan, Sangkwon JEONG, Youngkwon KIM, Kie-Joo CHO, and Seung-Hyub OH. "Experimental Investigation of Liquid Helium Pressurization Method for Liquid Propellant Rocket." JSME International Journal Series B 48, no. 2 (2005): 300–304. http://dx.doi.org/10.1299/jsmeb.48.300.

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Dron, M. M., and A. B. Yakovlev. "Analysis of properties of inertia regulator of thrust control system of liquid rocket engine." Omsk Scientific Bulletin. Series Aviation-Rocket and Power Engineering 5, no. 2 (2021): 98–105. http://dx.doi.org/10.25206/2588-0373-2021-5-2-98-105.

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The quality of the flight task of a space rocket system is determined among other things by the accuracy of maintaining and regulating the thrust of the rocket engine. Improving the accuracy and reducing errors in the engine mode control system will reduce the cost of space launches or allow you to put a large payload into orbit. The article presents a mathematical model of a controller with an inertial booster of a liquid-propellant rocket engine, identifies parameters and values that affect its accuracy, and considers measures to reduce static error

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Osipov, Viatcheslav V., Matthew J. Daigle, Cyrill B. Muratov, Michael Foygel, Vadim N. Smelyanskiy, and Michael D. Watson. "Dynamical Model of Rocket Propellant Loading with Liquid Hydrogen." Journal of Spacecraft and Rockets 48, no. 6 (November 2011): 987–98. http://dx.doi.org/10.2514/1.52587.

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Mohammadi, Kaveh, and Mahmoud Gorji. "Prediction of Amine-Based Liquid Rocket Propellant Shelf Life." Propellants, Explosives, Pyrotechnics 38, no. 4 (December 19, 2012): 541–46. http://dx.doi.org/10.1002/prep.201200091.

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A J Sriganapathy, Varunavan P, Ramana.S, Praveen Kumar R, and M.Ranjani. "Design and Analysis of Injector with Different Configurations for Liquid Rocket Engine." international journal of engineering technology and management sciences 7, no. 3 (2023): 353–76. http://dx.doi.org/10.46647/ijetms.2023.v07i03.046.

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The performance of the rocket is determined by atomization, mixing and combustion process. Injectors are used to control these processes. Suitable method should be followed, so that the performance should be increased. The injector implementation in rockets determines the performance of the nozzle that can be achieved. Injectors can be as simple as a number of small diameter holes arranged in carefully constructed patterns through which the fuel and oxidizer travel. The performance of an injector can be improved by either using a superior propellant combustion, increasing the mass flow rate, positioning the angle of the hole or by reducing the size & increasing the number of orifices on the injector plate. In here the position is varied for different angles for injectors to improve the performance of the rocket. The injector is modeled using the Solidworks and a two dimensional combustion analysis to be carried out using ANSYS Fluent for different cases.

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Yue, Chun Guo, Xin Long Chang, Shu Jun Yang, and You Hong Zhang. "Numerical Simulation of Interior Flow Field of a Variable Thrust Rocket Engine." Advanced Materials Research 186 (January 2011): 215–19. http://dx.doi.org/10.4028/www.scientific.net/amr.186.215.

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With the support of powerful calculation ability of computer and Fluent of CFD software, integrative simulation research of the variable thrust liquid propellant rocket engine was developed. Numerical simulation of interior flow field of a variable thrust rocket engine with flux-oriented injector was done. The distributions of pressure, temperature, molar fraction of product and flow mach numbers were attained. By the contrast of the calculation results, the effects of structure parameters and working condition etc. on total whole performance of variable thrust rocket engine were analyzed. The results also provided theoretic references for design and optimization of variable thrust rocket engine.

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Zhao, Na, Yong Gang Yu, and Yu Qiang Wang. "Numerical Simulation of the Spray Characteristics in Small Scale Liquid Rocket Engine Combustion Chamber." Advanced Materials Research 383-390 (November 2011): 7729–33. http://dx.doi.org/10.4028/www.scientific.net/amr.383-390.7729.

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The mathematical and physical model of the liquid propellant spray in straight nozzle was proposed for studying the performance characteristics of the small-scale liquid rocket engine. With the Fluent software, the numerical simulation was carried out. Sauter mean diameter (SMD) of the HAN-based liquid propellant (LP1846) in the engine combustor changing with spray pressure, nozzle diameter and the liquid surface tension were analyzed. The results indicate that: in the spray pressure region of 1.8MPa~3.0MPa, at a fixed spray pressure, the smaller is the nozzle diameter, the smaller is the droplets’ SMD and the relationship between the SMD and the nozzle diameter is approximately linearity; for the same nozzle diameter and spray pressure, the larger is the surface tension, the larger is the liquid droplets’ SMD.

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Pylypenko, O. V., O. D. Nikolayev, I. D. Bashliy, and O. M. Zavoloka. "Approach to numerical simulation of the spatial motions of a gas/liquid medium in a space stage propellant tank in microgravity with account for the hot zone." Technical mechanics 2022, no. 4 (December 15, 2022): 3–13. http://dx.doi.org/10.15407/itm2022.04.003.

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Space propulsion systems ensure multiple startups and shutdowns of the main liquid-propellant rocket engines in microgravity conditions for spacecraft preset motions and reorientation control. During the passive flight of a space stage (after its main engine shutdown), the liquid propellant in the tanks continues moving by inertia in microgravity and moves as far away from the propellant management device as possible. In this case, the pressurization gas is displaced to the propellant management device, which creates the potential danger of the gas entering the engine inlet in quantities unacceptable for multiple reliable engine restarts. In this regard, the determination of the parameters of fluid movement in propellant tanks under microgravity conditions is a pertinent problem to be solved in the designing of liquid-propellant propulsion systems. This paper presents an approach to the theoretical calculation of the parameters of motion of the gas–liquid system in the propellant tanks of today’s space stages in microgravity conditions. The approach is based on the use of the finite element method, the Volume of Fluid method, and up-to-date computer tools for finite-element analysis (Computer Aided Engineering - CAE systems). A mathematical simulation of the spatial motion of the liquid propellant and the formation of free gas inclusions in passive flight was performed, and the motion parameters and shape of the free liquid surface in the tank and the location of gas inclusions were determined. The liquid motion in a model spherical tank in microgravity conditions was simulated numerically with and without account for the hot zone near the tank head. The motion parameters of the gas-liquid interface in a model cylindrical tank found using the proposed approach are in satisfactory agreement with experimental data. The proposed approach will significantly reduce the extent of experimental testing of space stages under development.

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Yagodnikov, D. A., V. P. Aleksandrenkov, K. E. Kovalev, A. G. Grigoryants, and A. A. Drenin. "Study of Hydraulic Characteristics of the Cooling Path of a Model Liquid Rocket Engine Manufactured using Additive Technology of Selective Laser Melting." Herald of the Bauman Moscow State Technical University. Series Mechanical Engineering, no. 6 (129) (December 2019): 41–52. http://dx.doi.org/10.18698/0236-3941-2019-6-41-52.

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The article discusses hydraulic tests of a model combustion chamber of a liquid-propellant rocket engine with a cooling path made using additive selective laser melting technology. The values of the coefficient of hydraulic resistance in the range of Re = 10--2500 are obtained and the influence of the design features of the cooling tract and its manufacturing technology on the hydraulic characteristics is determined. The results of the performed hydraulic tests confirm the possibility of using additive technologies based on selective laser melting technology for the manufacture of fire and power walls of combustion chambers of liquid rocket engines.

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Ghaffar, Zulkifli Abdul, Ahmad Hussein Abdul Hamid, and Mohd Syazwan Firdaus Mat Rashid. "Spray Characteristics of Swirl Effervescent Injector in Rocket Application: A Review." Applied Mechanics and Materials 225 (November 2012): 423–28. http://dx.doi.org/10.4028/www.scientific.net/amm.225.423.

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Injector is one of the vital devices in liquid rocket engine (LRE) as small changes in its configurations and design can result in significantly different LRE performance. Characteristics of spray such as spray cone angle, breakup length and Sauter mean diameter (SMD) are examples of crucial parameters that play the important role in the performance of liquid propellant rocket engine. Wider spray cone angle is beneficial for widespread of fuel in the combustion chamber for fast quiet ignition and a shorter breakup length provides shorter combustion chamber to be utilized and small SMD will result in fast and clean combustion. There are several mechanisms of liquid atomization such as swirling, e.g. jet swirl atomization or introducing bubbles into the liquid and effervescent atomization. Introducing a swirl component in the flow can enhance the propellant atomization and mixing whereas introducing bubbling gas directly into the liquid stream inside the injector leads to finer sprays even at lower injection pressures. This paper reviews the influence of both operating conditions and injector internal geometries towards the spray characteristics of swirl effervescent injectors. Operating conditions reviewed are injection pressure and gas-to-liquid ratio (GLR), while the injector internal geometries reviewed are limited to swirler geometry, mixing chamber diameter (dc), mixing chamber length (lc), aeration hole diameter (da), discharge orifice diameter (do) and discharge orifice length (lo).

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Basharina, T. A., M. G. Goncharov, S. N. Lymich, V. S. Levin, and D. P. Shmatov. "Low-thrust liquid-propellant rocket engines as part of advanced ultralight rocket vehicle systems." Spacecrafts & Technologies 5, no. 1 (March 25, 2021): 5–13. http://dx.doi.org/10.26732/j.st.2021.1.01.

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This work examines the most promising design solutions for the creation of propulsion systems for ultra-light launch vehicles by small private enterprises in the rocket and space industry. Comparison of the metal consumption of the combustion chambers with the energy characteristics at different operating pressures showed that the most optimal operating pressure is 12,16 MPa. Comparison of the relative and absolute values of the masses of various configurations describes the nature of the relationship between the number of combustion chambers and the total mass of the propulsion system. It was found that nine-chamber propulsion systems with cameras made with extensive use of additive technologies best meet the key requirements. The analysis carried out includes an assessment of the design parameters of both various components and assemblies and the propulsion system as a whole. Various layouts of propulsion systems are considered in detail, the required degree of technological complexity of structures of various units and assemblies, their production cost are estimated. The ratio of the obtained mass-energy characteristics was achieved through the implementation of design solutions that became available due to the use of additive technologies. The obtained results of preliminary calculations demonstrate the applicability and efficiency of design solutions considered for use in the propelled propulsion system for a promising launch vehicle.

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Journal articles on the topic 'Liquid Propellant Rocket'

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