Keywords cloud
[Guidelines on elimination of large spillages of oxidizer NTO and fuel UDMH.
]]>5. Assessment of risk of toxic damage to people in case of a launch vehicle accident at flight
Organization:
Page: Kosm. teh. Raket. vooruž. 2024, (1); 40-50
DOI: https://doi.org/10.33136/stma2024.01.040
Language: English
Annotation: Despite stringent environmental requirements, modern launch vehicles/integrated launch vehicles (LV/ILV)
burn toxic propellants such as NTO and UDMH. Typically, such propellants are used in the LV/ILV upper
stages, where a small amount of propellant is contained; however, some LV/ILV still use such fuel in all
sustainer propulsion stages. For launch vehicles containing toxic rocket propellants, flight accidents may result
in the failed launch vehicle falling to the Earth’s surface, forming large zones of chemical damage to people
(the zones may exceed blast and fire zones). This is typical for accidents occurring in the first stage flight
segment, when an intact launch vehicle or its components (usually individual stages) with rocket propellants
will reach the Earth’s surface. An explosion and fire following such an impact will most likely lead to a massive
release of toxicant and contamination of the surface air. An accident during the flight segment of the LV/ILV
first stage with toxic rocket propellants, equipped with a flight termination system that implements emergency
engine shutdown in case of detection of an emergency situation, has been considered. To assess the risk
of toxic damage to a person located at a certain point, it is necessary to mathematically describe the zone
within which a potential impact of the failed LV/ILV will entail toxic damage to the person (the so-called zone
of dangerous impact of the failed LV/ILV). The complexity of this lies in the need to take into account the
characteristics of the atmosphere, primarily the wind. Using the zone of toxic damage to people during the fall
of the failed launch vehicle, which is proposed to be represented by a combination of two figures: a semicircle
and a half-ellipse, the corresponding zone of dangerous impact of the failed LV/ILV is constructed. Taking into
account the difficulties of writing the analytical expressions for these figures during the transition to the launch
coordinate system and further integration when identifying the risk, in practical calculations we propose to
approximate the zone of dangerous impact of the failed LV/ILV using a polygon. This allows using a known
procedure to identify risks. A generalization of the developed model for identifying the risk of toxic damage
to people involves taking into account various types of critical failures that can lead to the fall of the failed
LV/ILV, and blocking emergency engine shutdown during the initial flight phase. A zone dangerous for people
was constructed using the proposed model for the case of the failure of the Dnepr launch vehicle, where the
risks of toxic damage exceed the permissible level (10–6). The resulting danger zone significantly exceeds the
danger zone caused by the damaging effect of the blast wave. Directions for further improvement of the model
are shown, related to taking into account the real distribution of the toxicant in the atmosphere and a person’s
exposure to a certain toxic dose.
Key words: launch vehicle, critical failure, flight accident, zone of toxic damage to people, zone of dangerous impact of the failed launch vehicle, risk of toxic damage to people.
Bibliography:
- Hladkiy E. H. Protsedura otsenky poletnoy bezopasnosti raket-nositeley, ispolzuyuschaya geometricheskoe predstavlenie zony porazheniya obiekta v vide mnogougolnika. Kosmicheskaya technika. Raketnoe vooruzhenie: sb. nauch.-techn. st. Dnepropetrovsk: GP «KB «Yuzhnoye», 2015. Vyp. 3. S. 50 – 56. [Hladkyi E. Procedure for evaluation of flight safety of launch vehicles, which uses geometric representation of object lesion zone in the form of a polygon. Space Technology. Missile Weapons: Digest of Scientific Technical Papers. Dnipro: Yuzhnoye SDO, 2015. Issue 3. Р. 50 – 56. (in Russian)].
- Hladkiy E. H., Perlik V. I. Vybor interval vremeni blokirovki avariynogo vyklucheniya dvigatelya na nachalnom uchastke poleta pervoy stupeni. Kosmicheskaya technika. Raketnoe vooruzhenie: sb. nauch.-tech. st. Dnepropetrovsk: GP «KB «Yuzhnoye», 2011. Vyp. 2. s. 266 – 280. [Hladkyi E., Perlik V. Selection of time interval for blocking of emergency engine cut off in the initial flight leg of first stage. Space Technology. Missile Weapons: Digest of Scientific Technical Papers. Dnipro: Yuzhnoye SDO, 2011. Issue 2. Р. 266 – 280. (in Russian)].
- Hladkiy E. H., Perlik V. I. Matematicheskie modeli otsenki riska dlya nazemnykh obiektov pri puskakh raket-nositeley. Kosmicheskaya technika. Raketnoe vooruzhenie: sb. nauch.-techn. st. Dnepropetrovsk: GP «KB «Yuzhnoye», 2010. Vyp. 2. S. 3 – 19. [Hladkyi E., Perlik V. Mathematic models for evaluation of risk for ground objects during launches of launch-vehicles. Space Technology. Missile Weapons: Digest of Scientific Technical Papers. Dnipro: Yuzhnoye SDO, 2010. Issue 2. P. 3 – 19. (in Russian)].
- NPAOP 0.00-1.66-13. Pravila bezpeki pid chas povodzhennya z vybukhovymy materialamy promyslovogo pryznachennya. Nabrav chynnosti 13.08.2013. 184 s [Safety rules for handling explosive substances for industrial purposes. Consummated 13.08.2013. 184 p.
(in Ukranian)]. - AFSCPMAN 91-710 RangeSafetyUserRequirements. Vol. 1. 2016 [Internet resource]. Link : http://static.e-publishing.af.mil/production/1/afspc/publicating/
afspcman91-710v1/afspcman91-710. V. 1. pdf. - 14 CFR. Chapter III. Commercial space transportation, Federal aviation administration, Department of transportation, Subchapter C – Licensing, part 417 – Launch Safety, 2023 [Internet resource]. Link: http://law.cornell.edu/cfr/text/14/part-417.
- 14 CFR. Chapter III. Commercial space transportation, Federal aviation administration, Department of transportation, Subchapter C – Licensing, part 420 License to Operate a Launch Site. 2022 [Internet resource]. Link: http://law.cornell.edu/cfr/text/14/part-420.
- ISO 14620-1:2018 Space systems – Safety requirements. Part 1: System safety.
- 9 GOST 12.1.005-88. Systema standartov bezopasnosti truda. Obschie sanitarno-gigienicheskie trebovaniya k vozdukhu rabochei zony. [GOST 12.1.005-88. Labor safety standards system. General sanitary and hygienic requirements to air of working zone].
- 10 Rukovodyaschiy material po likvidatsii avarijnykh bolshykh prolivov okislitelya АТ (АК) i goruchego NDMG. L.:GIPKh, 1981, 172 s. [Guidelines on elimination of large spillages of oxidizer NTO and fuel UDMH. L.:GIPH, 1981, 172 p. (in Russian)].
- 11 Kolichestvennaya otsenka riska chimicheskykh avariy. Kolodkin V. M., Murin A. V., Petrov A. K., Gorskiy V. G. Pod red. Kolodkina V. M. Izhevsk: Izdatelskiy dom «Udmurtskiy universitet», 2001. 228 s. [Quantitative risk assessment of accident at chemical plant. Kolodkin V., Murin A., Petrov A., Gorskiy V. Edited by Kolodkin V. Izhevsk: Udmurtsk’s University. Publish house, 2001. 228 p. (in Russian)].
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Taking into consideration the high cost of neutralization units, which will be a factor hindering the wide-scale introduction of neutralization units to decrease technogenic load on environment of Ukraine, the option is proposed of reducing the costs during the use of thermal neutralization units by way of combining the function of oxidizer neutralization unit and fuel neutralization unit in a single universal neutralization unit.
]]>20. Studying the possibility of alternating delivery of rocket propellant wastes to a common thermal neutralization facility
Organization:
Yangel Yuzhnoye State Design Office, Dnipro, Ukraine
Page: Kosm. teh. Raket. vooruž. 2020, (1); 177-183
DOI: https://doi.org/10.33136/stma2020.01.177
Language: Russian
Annotation: The possibility is considered of using rocket propellant thermal neutralization units for decontamination of dangerous industrial wastes. The advantages of thermal neutralization units are considered, their operating principle is described, by the example of high-temperature toxic rocket propellants, the chemical reactions that take place in combustion chamber are formulated. The combustion chamber is a component of the neu tralization unit, it is in the combustion chamber, in the environment of created high temperatures, that the process of elimination of dangerous substances takes place. Taking into consideration the high cost of neutralization units, which will be a factor hindering the wide-scale introduction of neutralization units to decrease technogenic load on environment of Ukraine, the option is proposed of reducing the costs during the use of thermal neutralization units by way of combining the function of oxidizer neutralization unit and fuel neutralization unit in a single universal neutralization unit. The article substantiates the topicality and necessity of works to create the universal thermal neutralization unit from the viewpoint of economic and ecological aspects. The article presents a generalized description of technology and methodology of research tests of pilot samples of assemblies for high-temperature rocket propellants vapor and industrial wastewater supply into the neutralization unit. The assemblies for high-temperature rocket propellants vapor and industrial wastewater supply are considered as most critical components of the universal neutralization unit from the viewpoint of neutralized substance changing. The experiments were conducted on water solutions of rocket propellants that in this case simulated the contact of internal cavities of supply assemblies with aggressive toxic media. The conditions were created at which the probability existed of interaction of rocket propellants residues in stagnation zones at the moment of changing the supplied propellant component. In the frameworks of research tests of pilot samples, the obtained results were considered and analyzed. The findings are presented that confirm practical feasibility of using integrated supply assemblies.
Key words: neutralization unit, supply assemblies, alternate supply, rocket propellants interaction, universal thermal neutralization unit
Bibliography:
1. Kolesnikov S. V. Okislenie nesimmetrichnogo dimetilgidrazina (geptila) i identifikatsiia produktov ego prevrashcheniia pri prolivakh. Monografiia. NP “SibAK”, Novosibirsk, 2014.
2. Zhidkoe raketnoe toplivo v regione OBSE: obzor aspektov utilizatsii. FSC.DEL/443/07/Rev. 2. 23 okt. 2008 g.
3. Egorychev V. S., Kondrusev V. S. Topliva khimicheskikh raketnykh dvigatelei. Samara, 2007.
4. Kasimov А. М., Semenov V. Т., Shcherban’ N. H., Miasoedov V. V. Sovremennye problemy i resheniia v sisteme upravleniia opasnymi otkhodami. Kharkiv, 2008.
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Keywords cloud
Moreover, they are launched in the lower stratosphere or in the troposphere so that there is no need to place oxidizer supply on board.
]]>17. Peculiarities of Dynamics of Recoverable Part of Stage of Aircraft-Type Configuration with Turbojet Engine
Organization:
Yangel Yuzhnoye State Design Office, Dnipro, Ukraine
Page: Kosm. teh. Raket. vooruž. 2018 (2); 143-150
DOI: https://doi.org/10.33136/stma2018.02.143
Language: Russian
Annotation: Basic dynamic properties of the reentry part of the aircraft-type first stage were examined when turbojet engine is used in the recovery phase. Such configuration can be of interest because turbojets have considerably smaller rate of flow in comparison to rocket engines. Moreover, they are launched in the lower stratosphere or in the troposphere so that there is no need to place oxidizer supply on board. This recovery plan differs from an alternative rocket recovery system and, from our point of view, provides more efficient usage of the fuel stores because it doesn’t require the main propulsion to be started in the recovery phase. Besides the analysis of qualitative characteristics of the descend phase for this stage, the efficiency of a wing with moderate values of maximum aerodynamic characteristics and a turbojet was studied. In this case three ways for stage recovery were investigated. The first one implied unguided descend with zero angle of attack assuming that the stage is statically stable. This descend trajectory was considered as standard and was used to evaluate the efficiency of the wing and turbojet with relatively small propulsion. The second and the third design cases offered the gliding guided descend with turbojet being started only in the lower stratosphere. The last two cases used the same program for the angle of attack. The possibility to ensure permissible overload values at the critical points of the descend trajectory and acceptable values of kinematic characteristics at the earth surface tangency point are also of great interest. Thereby the program for the angle of attack was developed in a way that allowed kinematic characteristics on touchdown be as close as possible to the corresponding values, shown by civil and/or military-transport heavy aircraft. Simulation was conducted on Microsoft Visual Studio 2010.
Key words: guided descent, turbojet, kinematic characteristics, tangency point, civil aviation
Bibliography:
1. Kuznetsov Y. L., Ukraintsev D. S. Analysis of Impact of Flight Scheme of Stage with Rocket-Dynamic Recovery System on Payload Capability of Medium-Class Two-Stage Launch Vehicle. New of S. P. Korolev Samara State Aerospace University (National Research University). 2016. Vol. 15, No. 1. P. 73-80. https://doi.org/10.18287/2412-7329-2016-15-1-73-80
2. Andreyevsky V. V. Spacecraft Earth Descent Dynamics М., 1970. 230 p.
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Keywords cloud
RD868 engine oxidizer and fuel pumps;
]]>9. The Impact of Worm Design on Power and Anti-Cavitation Properties of Worm-Centrifugal Pumps
Organization:
Yangel Yuzhnoye State Design Office, Dnipro, Ukraine
Page: Kosm. teh. Raket. vooruž. 2018 (2); 76-82
DOI: https://doi.org/10.33136/stma2018.02.076
Language: Russian
Annotation: In the present-day rocket engineering, the liquid rocket engines with pump feed system have gained wide acceptance. As a rule, the pumps used in liquid rocket engines are screw-axifugal. The screw serves to increase pressure upstream of the axifugal wheel, thus ensuring its cavitation-free operation. The screws used in the screw-axifugal pumps of liquid rocket engines may be of two types: with constant and variable step. The screws with constant step are easier to calculate, profile and manufacture as compared to the screws with variable step. As known from the literature, the use of the screw with variable step increases power characteristics of the screw-axifugal pump. The purpose of investigation is comparative analysis of cavitation and power characteristics of the following high-speed low-consumption screw-axifugal pumps of liquid rocket engines with jointed screws, screws of constant and variable step: RD868 engine oxidizer and fuel pumps; RD859 engine fuel pumps; RD861K engine fuel pumps. Besides, the analysis has been made of the impact of design features and geometrical dimensions of the screws with variable and constant step on power characteristics of the screw-axifugal high-speed lowconsumption pumps of liquid rocket engines. Special attention has been given to the analysis of anticavitation properties of the pumps with screws of variable step and pumps with jointed screws. Based on the results of investigation, it has been ascertained that when using the joint screws and screws with variable step instead of the screws with constant step in the high-speed low-consumption screw-axifugal pumps of liquid rocket engines, the pump delivery head increases from 0.65 to 3.83%, the efficiency increases up to 1.7%. The use of jointed screw and screw of variable step as compared with the screw of constant step does not have any impact on cavitation properties of low-consumption crew-axifugal pumps of liquid rocket engines.
Key words: pressure characteristic, cavitation characteristic, inducers of the variable-pitch, continuous-pitch inducers, pump efficiency
Bibliography:
1. Pre-burner operating method for rocket turbopump: Patent 6505463 USA: MPK F02K9/48 / William D. Kruse, Thomas J. Mueller, John J. Weede (USA); Northrop Grumman Corporation. No. 20020148215; claimed 17.01.2001; published 14.01.2003, Bulletin No. 09/761,957. 5 p.
2. Hybrid rocket motor using a turbopump to pressurize a liquid propellant constituent: Patent 6640536 USA: MPK F02K9/50, F02K9/48, F02K9/46, F02K9/72, F02K9/56 / Korey R. Kline, Kevin W. Smith, Eric E. Schmidt, Thomas O. Bales; Hy Pat Corporation (Miami, FL). No. 20030136111; claimed 22.01.2002; published 04.11.2003, Bulletin No. 10/054,646. – 11 p.
3. Chebayevsky V. F., Petrov V. I. Cavitation Characteristics of High-Speed Auger-Centrifugal Pumps. М., 1973. 152 p.
4. Petrov V. I., Chebayevsky V. F. Cavitation on High-Speed Impeller Pumps. М., 1982. 192 p.
5. Ovsyanikov V. B., Borovsky B. I. Theory and Calculation of Liquid Rocket Engines Generator Sets. М, 1986. 376 p.
6. Borovsky B. I. Power Parameters and Characteristics of High-Speed Impeller Pumps. М., 1989. 181 p.
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Keywords cloud
Oxidizer Feedline Structural Optimization Results Authors: Veskov E. (2017) "Oxidizer Feedline Structural Optimization Results" Космическая техника. "Oxidizer Feedline Structural Optimization Results" Космическая техника. quot;Oxidizer Feedline Structural Optimization Results", Космическая техника. Oxidizer Feedline Structural Optimization Results Автори: Veskov E. Oxidizer Feedline Structural Optimization Results Автори: Veskov E. Oxidizer Feedline Structural Optimization Results Автори: Veskov E. Oxidizer Feedline Structural Optimization Results Автори: Veskov E.
]]>15. Oxidizer Feedline Structural Optimization Results
Organization:
Yangel Yuzhnoye State Design Office, Dnipro, Ukraine
Page: Kosm. teh. Raket. vooruž. 2017 (2); 77-82
Language: Russian
Annotation: Two design options of manifold and dividing valve are considered, the loss calculation by analytical and numerical methods has been made. Based on the calculation results, the optimal design option has been selected. The calculation correctness is confirmed as a result of development tests of the design.
Key words:
Bibliography:
1. Idel’chik I. E. Guide on Hydraulic Resistances / Under the editorship of M. O. Steinberg. 3rd edition revised and enlarged. М., 1992. 672 p.
2. Yan’shin B. I. Hydrodynamic Characteristics of Regulating Valves and Pipeline Elements. М., 1965. 259 p.
3. Gurevich D. F. Calculation and Designing of Pipeline Fittings: Calculation of Pipeline Fittings. 5th edition. М., 2008. 480 p.
4. Frenkel N. Z. Hydraulics. М., L., 1956. 451 p.
5. Reference Book on Hydraulics, Hydraulic Machines, and Hydraulic Actuators / Under the editorship of B. B. Nekrasov. Minsk, 1985.
6. Alyamovsky A. A. “Solid Works” Computer Modeling in Engineering Practice. Saint Petersburg, 2012. 445 p.
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| Singapore | Singapore; Singapore; Singapore; Singapore; Singapore; Singapore; Singapore | 7 |
| Canada | Toronto; Toronto; Toronto; Monreale | 4 |
| Unknown | ; Hong Kong | 2 |
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| Romania | Voluntari | 1 |
2017 (2); 65-69 Language: Russian Annotation: To support ignition of Taurus LV first stage core structure engine, an oxidizer dividing valve has been developed that ensures minimal hydraulic resistance, opening time and hydraulic impact.
]]>13. Experience in Development of Isolating Fuel Valve with Pneumatic Drive and Hydraulic Brake for Operation under Cryogenic Conditions
Organization:
Yangel Yuzhnoye State Design Office, Dnipro, Ukraine
Page: Kosm. teh. Raket. vooruž. 2017 (2); 65-69
Language: Russian
Annotation: To support ignition of Taurus LV first stage core structure engine, an oxidizer dividing valve has been developed that ensures minimal hydraulic resistance, opening time and hydraulic impact. The paper considers the valve design, main phases of its ground development testing and basic critical comments on the design made in the process of manufacturing and testing.
Key words:
Bibliography:
1. Report on the Results of Isolation Valve Check Tests Taurus-II.21.17039.203ОТ / Yuzhnoye SDO. Dnepropetrovsk, 2011. 30 p.
2. Report on the Results of Isolation Valve Check Tests Taurus-II.21.17050.203ОТ / Yuzhnoye SDO. Dnepropetrovsk, 2011. 23 p.
3. Report-Conclusion on the Results of Isolation Valve Developmental Tests 2TRS2S1.94.7204.0000.0000.00.0 ОЗ / Yuzhnoye SDO. Dnepropetrovsk, 2011. 161 p.
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| Singapore | Singapore; Singapore; Singapore; Singapore; Singapore; Singapore; Singapore; Singapore | 8 |
| Canada | Toronto; Toronto; Toronto; Toronto; Toronto; Monreale | 6 |
| China | Pekin; Pekin | 2 |
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Key words: oxidizer , fuel , saturation by helium , tank container , transportation Bibliography: 1. oxidizer , fuel , saturation by helium , tank container , transportation .
]]>6. Investigation into Peculiarities of Delivery to Launch Base of Rocket Propellant with Specified Gasing
Organization:
Yangel Yuzhnoye State Design Office, Dnipro, Ukraine
Page: Kosm. teh. Raket. vooruž. 2019, (1); 38-44
DOI: https://doi.org/10.33136/stma2019.01.038
Language: Russian
Annotation: This article considers the issue of achievement of the specified value of propellants saturation by helium after their delivery from the manufacturers to the launch site. Knowing the fact that propellants gas saturation or gas separation processes are labour-consuming and costly this issue is of immediate interest. In order to solve this problem number of factors have been considered, which determine the value of gas saturation in the propellants delivered to the launch site and procedure to control the value of gas saturation by the fuel manufacturer has been developed. This procedure implies that shipping tank container is pressurized after being fueled with propellants at the manufacturer’s, the pressure is characterized by the value of the known initial deficit or excess of gas in the propellants, following which tank container is delivered to the launch site. During transportation tank container is subjected to various kinds of mechanical actions (vibration, rolling and pitching in the sea, braking, transshipment), therefore intensive mixing of propellants occur. As propellants mix, process of propellant saturation occurs when certain amount of gas transits from tank container’s gas volume into the liquid, therefore certain gas saturation is reached. Article includes the measuring results of the gas liquid medium parameters inside the tank containers with fuel in the process of fuel transportation to Ukraine from PRC factories and estimations of the measuring results using the developed model which confirmed the quantitative nature of the mass exchange processes, included in the model, going on in the gas liquid medium during transportation of the tank container with fuel equipment. It has been determined that due to inevitable errors in the measuring of the specified parameters by the tank container, the achievement of the specified gas saturation with high precision is problematic. In spite of the fact that this procedure does not provide exact value of the specified gas saturation, its application will accelerate and make cheaper the process of fuel preparation for filling operations at the launch site, which is especially relevant in case of fuel saturation by helium. Based on this fuel saturation by helium procedure, the complex technology is suggested, providing controlled gas saturation during fuel delivery and subsequent adjustment of gas saturation using launch site equipment. Therefore, this article develops and studies the original model of the controlled gas saturation of the fuel during its delivery to the consumer. Alternative of the practical use of the study results is suggested in the form of the complex technology of fuel saturation by helium, delivered in the tank containers from the manufacturer to the launch site.
Key words: oxidizer, fuel, saturation by helium, tank container, transportation
Bibliography:
1. Volskiy A. P. Kosmodrom. M.: Voenizdat, 1977. 311 p.
2. Stepanov A. N., Vorobiev A. M., Grankin B. K. Kompleksy zapravki raket I kosmicheskikh apparatov. SPB:OM-PRESS, 2004. 26 p.
3. Kiriyanova A. N., Matveeva O. P. Opredelenie kolebania davlenia v gazovoy polosti hermetychikh emkostey transportnozapravochnykh containerov dlya raketnykh topliv pri temperaturnykh vozdeistviyakh/ Nauka i innovatsii. 2016. Vyp. 7.
4. Berezhkovskiy M. I. Khranenie i transportirovka khimicheskykh produktov. – M.: Khimia, 1973. – 272 s.
5. Perepelkin K. Ye., Matveev V. S. Gazovye emulsii. L.:Khimia, 1979. 200 p.
6. Issledovanie protsessov degazirovaniya komponentov topliva v conteinere-tsisterne pri dostavke topliva potrebitelyu. Cyclone4M 21.18425.174 OT: Techn. report. Dnepropetrovsk: Yuzhnoye SDO, 2017. 39 p.
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Keywords cloud
Different combinations of oxidizer and fuel are considered based on which the power characteristics of hybrid rocket motors were determined.
]]>12. Increasing Effectiveness of Meteorological Rockets when Using Hybrid Motors
Organization:
Yangel Yuzhnoye State Design Office, Dnipro, Ukraine
Page: Kosm. teh. Raket. vooruž. 2016 (1); 76-81
Language: Russian
Annotation: A comparative assessment of effectiveness of usage of hybrid rocket motors and solid rocket motors for meteorological rocket with propellant mass of 540 kg is made. Different combinations of oxidizer and fuel are considered based on which the power characteristics of hybrid rocket motors were determined. Based on the results of investigation, an assessment is given of ballistic effectiveness of hybrid rocket motors, the peculiarities of combustion process and of selection and designing of solid propellant charge shape are analyzed.
Key words:
Bibliography:
Full text (PDF) || Content 2016 (1)
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The article considers the preparation of propellant for filling the launch vehicle stages with high-boiling propellant components of nitrogen tetroxide (oxidizer) and unsymmetrical dimethyl hydrazine (fuel) in terms of propellant saturation with helium and denitrogenation.
]]>5. Fueling-neutralization stations. New developments and applications
Organization:
Yangel Yuzhnoye State Design Office, Dnipro, Ukraine
Page: Kosm. teh. Raket. vooruž. 2023 (1); 48-55
DOI: https://doi.org/10.33136/stma2023.01.048
Language: Ukrainian
Annotation: The article dwells on development and study of the multifunctional operations while preparing propellant components for launch vehicle tank filling by high-boiling propellant components at the neutralization stations. The article considers the preparation of propellant for filling the launch vehicle stages with high-boiling propellant components of nitrogen tetroxide (oxidizer) and unsymmetrical dimethyl hydrazine (fuel) in terms of propellant saturation with helium and denitrogenation. Usually these issues refer to the common technology of propellant preparation and are tackled sequentially: first, propellant is drained into the filling tank of the filling system, then the propellant is denitrogened, for example purging the propellant by helium under atmospheric pressure in the tank, then propellant is saturated with helium to the given concentration by bubbling helium in the propellant, maintaining the set pressure of helium in the tank. This technology significantly complicates the process of propellant preparation, increases helium consumption, as well as the amount of the generated vapor, which requires recycling in the neutralization units.
This article studies multifunctional operations, where the propellant is simultaneously drained from delivery vehicles, saturated with helium and denitrogenated. Amount of residual nitrogen in the propellant and the main direction of deep denitrogenation of the propellant are calculated. Amount of generated vapor and consumed helium are determined. The process of propellant draining by extrusion, maintaining the given pressure in the tank and alternating the propellant drain on the closed vent device (compression) and open vent device (decompression) is studied.
As a result, the theoretical justification of multifunctional operations in preparation of high–boiling propellant components to fill the launch-vehicle stages is presented.
Key words: saturation with helium and denitrogenation of the propellant, drain with closed vent device, excessive pressure draining, gas-vapor mixture, neutralization system
Bibliography:
1. Pozdeev G. L., Kucherenko R. A., Kucherenko T. V. Issledovanie osobennostey dostavki na kosmodrom komponentov raketnogo topliva s zadannym gazonasyschenniem. Kosmicheskaya technika. Raketnoe vooruzhenie: sb. nauch.-techn. st. / GP «KB «Yuzhnoye». Dnepr, 2019. Vyp. 1. S. 38–44. https://doi.org/10.33136/stma2019.01.038
2. Pozdeev G. L. Razrabotka I issledovanie metodov obespecheniya zadannyh parametrov pri zapravke: dis. na soiskanie kand. tehn. nauk. GP «KB «Yuzhnoye», 1978. Dnepropetrovsk. 123 s.
3. KRK «Tsiklon-4M». Zapravochno-neitralizatsionnaya stantsia. Tehnicheskiy proekt. С4М YZH-ANL 02802. GP «KB «Yuzhnoye». 2017. 108 s.
4. KRK «Tsiklon-4M». Predlozheniya po helirovaniyu komponentov topliva dlya zapravki 2-oy stupeni rakety-nositelya na ZNS s uchetom osobennostey dostavki topliva v hermetichnyh konteynerah-tsisternah. Nauchno-tehnicheskiy otchet. Tsiklon-4M. 21.18668.174 OT. GP «KB «Yuzhnoye», 2019. 33 s.
5. Sposob zakrytoy zapravki toplivnogo baka zhidkim toplivom i systema dlya ego osuschestvlenniya: pat. RU 2489327. В64F1/28.
Full text (PDF) || Content 2023 (1)
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Keywords cloud