Search Results for “thrust” – Collected book of scientific-technical articles https://journal.yuzhnoye.com Space technology. Missile armaments Tue, 05 Nov 2024 21:15:15 +0000 en-GB hourly 1 https://journal.yuzhnoye.com/wp-content/uploads/2020/11/logo_1.svg Search Results for “thrust” – Collected book of scientific-technical articles https://journal.yuzhnoye.com 32 32 2.1.2020 Analysis of development trends of design parameters and basic characteristics of missiles for the advanced multiple launch rocket systems https://journal.yuzhnoye.com/content_2020_1-en/annot_2_1_2020-en/ https://journal.yuzhnoye.com/?page_id=31001
The mathematical model of the guided missile provides adequate accuracy for design study to determine depending on the main design parameters: overall dimensions and mass characteristics of the guided missile in general and its structural comp onents and subsystems; power, thrust and consumption characteristics of the rocket motor; aerodynamic and ballistic characteristics of the guided missile.
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2. Analysis of development trends of design parameters and basic characteristics of missiles for the advanced multiple launch rocket systems

Authors:

Aksenenko A. V.1, Hurskyi O. I.1, Klochkov A. S.1, Kondratiuk Ye. A.1, Senkin V. S.2, Siutkina-Doronina S. V.2

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine1; The Institute of Technical Mechanics, Dnipro, Ukraine2

Page: Kosm. teh. Raket. vooruž. 2020, (1); 13-25

DOI: https://doi.org/10.33136/stma2020.01.013

Language: Russian

Annotation: The scientific and methodological propositions for the designing single-stage guided missiles with the solid rocket motors for advanced multiple launch rocket systems are defined. The guided missiles of multiple launch rocket system are intended for delivering munitions to the given spatial point with required and specified kinematic motion parameters at the end of flight. The aim of the article is an analysis of the development trends of the guided missiles with the solid rocket motors for the multiple launch rocket systems, identifying the characteristics and requirements for the flight trajectories, design parameters, control programs, overall dimensions and mass characteristics, structural layout and aerodynamic schemes of missiles. The formalization of the complex task to optimize design parameters, trajectory parameters and motion control programs for the guided missiles capable of flying along the ballistic, aeroballistic or combined trajectories is given. The complex task belongs to a problem of the optimal control theory with limitations in form of equa lity, inequality and differential constraints. To simplify the problem, an approach to program forming is proposed for motion control in the form of polynomial that brings the problem of the optimal control theory to a simpler problem of nonlinear mathematical programming. When trajectory parameters were calculated the missile was regarded as a material point of variable mass and the combined equations for center-of-mass motion of the guided missile with projections on axes of the terrestrial reference system were used. The structure of the mathematical model was given along with the calculation sequence of the criterion function that was used for determination of the optimal parameters, programs and characteristics. The mathematical model of the guided missile provides adequate accuracy for design study to determine depending on the main design parameters: overall dimensions and mass characteristics of the guided missile in general and its structural comp onents and subsystems; power, thrust and consumption characteristics of the rocket motor; aerodynamic and ballistic characteristics of the guided missile. The developed methodology was tested by determining design and trajectory parameters, overall dimensions and mass characteristics, power and ballistic characteristics of two guided missiles with wings for advanced multiple launch rocket systems produced by the People’s Republic of China, using the limited amount of information available in the product catalog.

Key words: multiple launch rocket systems (MLRS), complex problem of the optimal control theory, problem of nonlinear mathematical programming, main solid rocket motor, limitations for motion parameters and basic characteristics of the guided missiles

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2.1.2020 Analysis of development trends of design parameters and basic characteristics of missiles for the advanced multiple launch rocket systems
2.1.2020 Analysis of development trends of design parameters and basic characteristics of missiles for the advanced multiple launch rocket systems
2.1.2020 Analysis of development trends of design parameters and basic characteristics of missiles for the advanced multiple launch rocket systems

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]]> 8.1.2024 Theoretic-experimental evaluation of the solid-propellant grain erosive burning https://journal.yuzhnoye.com/content_2024_1-en/annot_8_1_2024-en/ Mon, 17 Jun 2024 08:41:58 +0000 https://journal.yuzhnoye.com/?page_id=35027
2024, (1); 72-77 DOI: https://doi.org/10.33136/stma2024.01.072 Language: Ukrainian Annotation: The high demands for the flow rate and thrust characteristics specified for the modern solid-propellant rocket motors (SRM) under the strict mass and overall dimensions constraints require high level of mass fraction of propellant. This may play both negative (off-design increase of chamber pressure) and positive role (for example, increasing the launch thrust-to-weight ratio of the rocket).
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8. Theoretic-experimental evaluation of the solid-propellant grain erosive burning

Автори: Taran M. V., Moroz V. G.

Organization: Yangel Yuzhnoye State Design Office, Dnipro, Ukraine

Page: Kosm. teh. Raket. vooruž. 2024, (1); 72-77

DOI: https://doi.org/10.33136/stma2024.01.072

Language: Ukrainian

Annotation: The high demands for the flow rate and thrust characteristics specified for the modern solid-propellant rocket motors (SRM) under the strict mass and overall dimensions constraints require high level of mass fraction of propellant. And in the process of propellant grain combustion, erosive burning often takes place (increase of propellant burning rate depending on combustion products flow rate along the grain channel). This may play both negative (off-design increase of chamber pressure) and positive role (for example, increasing the launch thrust-to-weight ratio of the rocket). It is typical of the main SRMs of various rocket systems (multiple launch rocket systems, anti-aircraft guided missiles, tactical missiles, boosters). This paper proposes a methodology for calculating the internal ballistic characteristics of a solid propellant rocket motor under erosive burning, which is relatively time and resource consuming. The methodology is based on equidistant model of propellant grain combustion, where grain is divided lengthwise into a number of intervals. For any point of time during the engine operation, burning area and port area of each interval are calculated, taking into account erosive impact on each interval; total burning area is the sum of all intervals burning areas. Gas flow rate in each interval of the grain channel is calculated using gas-dynamic equations. The motor mass flow rate is a mass input sum of all the intervals; and the burning rate in each interval is estimated with proper erosion factor. The combustion chamber pressure had been calculated for four erosive burning models proposed by different authors. All the models showed convergence with the experimental SRM test data sufficient for engineering estimate (in particular, for maximum chamber pressure and combustion time). Selected as a result erosive burning model may be used to design new motors with solid propellants similar in chemical composition, and the model parameters are to be further customized using the test specimens.

Key words: rocket motor, solid propellant, erosive burning, internal ballistic characteristics

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  8. William Orvis. EXCEL dlya uchenykh, inzhenerov i studentov. Kiev: «Junior», 1999.

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8.1.2024 Theoretic-experimental evaluation of the solid-propellant grain erosive burning
8.1.2024 Theoretic-experimental evaluation of the solid-propellant grain erosive burning
8.1.2024 Theoretic-experimental evaluation of the solid-propellant grain erosive burning

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]]> 7.1.2024 Selection of the functional units for the Cyclone-4M ILV separation system https://journal.yuzhnoye.com/content_2024_1-en/annot_7_1_2024-en/ Fri, 14 Jun 2024 11:36:31 +0000 https://journal.yuzhnoye.com/?page_id=34957
Brief characteristics of these systems are given, based on the gas-reactive nozzle thrust, braking with solid-propellant rocket engines, separating with spring or pneumatic pushers.
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7. Selection of the functional units for the Cyclone-4M ILV separation system

Authors:

Logvinenko A. I., Makarenko A. O., Khomyak V. O., Olesiyuk A. A.

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine

Page: Kosm. teh. Raket. vooruž. 2024, (1); 61-71

DOI: https://doi.org/10.33136/stma2024.01.061

Language: Ukrainian

Annotation: Separation of the spent LV stages is one of the important problems of the rocket technology, which requires the comprehensive analysis of different types of systems, evaluation of their parameters and structural layouts. Basic requirements are specified that need to be taken into account when engineering the separation system: reliable and safe separation, minimal losses in payload capability, keeping sufficient distance between the stages at the moment of the propulsion system start. Detailed classification of their types («cold», «warm», «hot», «cold-launched» separation) is given and their technical substance with advantages and drawbacks is described. Certain types of «cold» and «warm» separation of the spent stages of such rockets as Dnepr, Zenit, Antares, Falcon-9 with different operating principle are introduced – braking with the spent stage and pushing apart two stages. Brief characteristics of these systems are given, based on the gas-reactive nozzle thrust, braking with solid-propellant rocket engines, separating with spring or pneumatic pushers. Development of the separation system for the advanced Cyclone-4M ILV is taken as an example and design sequence of stage separation is suggested: determination of the necessary separation velocity and capability of the separation units, determination of the number of active units, calculation of design and energy parameters of the separation units, analysis of the obtained results followed by the selection of the separation system. Use of empirical dependences is shown, based on the great scope of experimental and theoretical activities in the process of design, functional testing and flight operation of similar systems in such rockets as Cyclone, Dnepr and Zenit. According to the comparative analysis results, pneumatic separation system to separate Cyclone-4M Stages 1 and 2 was selected as the most effective one. Its basic characteristics, composition, overall view and configuration are specified. Stated materials are of methodological nature and can be used to engineer the separation systems for LV stages, payload fairings, spacecraft etc.

Key words: separation system, functional units of separation, «cold separation», «warm separation», pneumatic pusher, spring pusher, SPRE, gas-reactive nozzles, Zenit LV, Dnepr LV, Falcon 9 rocket, Cyclone-4М LV.

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7.1.2024 Selection of the functional units for the Cyclone-4M ILV separation system
7.1.2024 Selection of the functional units for the Cyclone-4M ILV separation system
7.1.2024 Selection of the functional units for the Cyclone-4M ILV separation system

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]]> 2.1.2024 New and advanced liquid rocket engines of the Yuzhnoye SDO https://journal.yuzhnoye.com/content_2024_1-en/annot_2_1_2024-en/ Wed, 12 Jun 2024 15:04:41 +0000 https://journal.yuzhnoye.com/?page_id=34964
Significant steps in this area have been made by the experts from the Yuzhnoye propulsion division and then subsequent manufacture and delivery by Yuzhmash PA of the engine intended for the European rocket Vega Stage 4; and designing the individual components for the engines with thrusts ranging from 500 kgf to 200 tf ordered by foreign customers. This article provides the review of current and scheduled activities of the Yuzhnoye SDO to develop the liquid rocket engines within the thrust ranges from ~ 40 kgf to ~ 500 tf. Key words: LOX-kerosene liquid rocket engines , hypergolic propellant liquid rocket engines , staged combustion cycle , main rocket engine , thrust , specific thrust impulse. LOX-kerosene liquid rocket engines , hypergolic propellant liquid rocket engines , staged combustion cycle , main rocket engine , thrust , specific thrust impulse.
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2. New and advanced liquid rocket engines of the Yuzhnoye SDO

Authors:

Prokopchuk O. O., Shulga V. A.

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine

Page: Kosm. teh. Raket. vooruž. 2024, (1); 9-18

DOI: https://doi.org/10.33136/stma2024.01.009

Language: Ukrainian

Annotation: Specialized design office for liquid engines was established on July 22, 1958 to develop engines and propulsion systems, powered by liquid propellants to be installed on the combat missile systems and integrated launch vehicles (LV), developed by Yuzhnoye SDO. Moreover, liquid engines design office was assigned with manufacturing and testing of the main rocket engines, developed by NPO Energomash and to be installed on Yuzhnoye-developed launch vehicles. Over the past 66 years Yuzhnoye SDO has developed more than 40 liquid rocket engines (LRE) of various purpose, designed both to gas-generator cycle and to staged combustion cycle. Seventeen of them were commercially produced by Yuzhmash PA and installed on launch vehicles. Nowadays Yuzhnoye propulsion experts keep working on development of the advanced liquid rocket engines powered both by cryogenic and hypergolic propellants, which satisfy the majority of launch service market demands. Within the framework of extensive cooperation with foreign space companies, on a contract basis, Yuzhnoye propulsion experts are working on the design and development testing of the liquid rocket engines, as well as their components. The accumulated vast experience in the development of liquid rocket engines nowadays enables high scientific and technical level in the creation of up-to-date engines, demanded in the world market. Significant steps in this area have been made by the experts from the Yuzhnoye propulsion division and then subsequent manufacture and delivery by Yuzhmash PA of the engine intended for the European rocket Vega Stage 4; and designing the individual components for the engines with thrusts ranging from 500 kgf to 200 tf ordered by foreign customers. This article provides the review of current and scheduled activities of the Yuzhnoye SDO to develop the liquid rocket engines within the thrust ranges from ~ 40 kgf to ~ 500 tf.

Key words: LOX-kerosene liquid rocket engines, hypergolic propellant liquid rocket engines, staged combustion cycle, main rocket engine, thrust, specific thrust impulse.

Bibliography:
  1. Zhidkostnye raketnye dvigateli, dvigatelnye ustanovki, bortovye istochniki moschnosti, razrabotannye KB dvigatelnykh ustanovok GP«KB «Yuzhnoye». Za nauk. red. akad. NAN Ukrainy S.M. Konyukhova, kand. tekhn. nauk V.M. Shnyakina. Dnipropetrovsk: DP «KB «Pivdenne», 2008. 466 ark.
  2. Prokopchyuk O. O., Shulga V. A., Khromyuk D. S., Sintyuk V. O. Zhidkostnye raketnye dvigateli GP«KB «Yuzhnoye»: nauk.-tekhn. zbirnyk. Za nauk. red. akademika NAN Ukrainy
    O. V. Degtyareva. Dnipro: ART-PRES, 2019. 440 ark.

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2.1.2024 New and advanced liquid rocket engines of the Yuzhnoye SDO
2.1.2024 New and advanced liquid rocket engines of the Yuzhnoye SDO
2.1.2024 New and advanced liquid rocket engines of the Yuzhnoye SDO

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]]> 10.1.2020 Calculation and selection of parameters for a propellant consumption diagram of dual-thrust main SRM https://journal.yuzhnoye.com/content_2020_1-en/annot_10_1_2020-en/ https://journal.yuzhnoye.com/?page_id=31037
Calculation and selection of parameters for a propellant consumption diagram of dual-thrust main SRM Authors: Yenotov V. The equations obtained show that the required parameters of diagrams (including consumption-thrust characteristics difference between the starting and cruise modes) can be ensured due to varying either case diameter or propellant combustion rate or due to combined variation of these values. (2020) " Calculation and selection of parameters for a propellant consumption diagram of dual-thrust main SRM" Космическая техника. " Calculation and selection of parameters for a propellant consumption diagram of dual-thrust main SRM" Космическая техника. quot; Calculation and selection of parameters for a propellant consumption diagram of dual-thrust main SRM", Космическая техника. Calculation and selection of parameters for a propellant consumption diagram of dual-thrust main SRM Автори: Yenotov V.
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10. Calculation and selection of parameters for a propellant consumption diagram of dual-thrust main SRM

Authors:

Yenotov V. H., Pustovharova O. V.

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine

Page: Kosm. teh. Raket. vooruž. 2020, (1); 99-106

DOI: https://doi.org/10.33136/stma2020.01.099

Language: Russian

Annotation: The main solid rocket motors of surface-to-air missiles and some short-range missiles have, as a rule, two operation modes – starting (augmented rating) and cruise (with decreased propellant consumption level). The methods to calculate intraballistic characteristics of such motors have a number of peculiarities, which set them apart from the methods of determining the characteristics of motors with constant propellant consumption level. The purpose of this article is to analyze such peculiarities, design methods, to find interrelation between the parameters of propellant consumption diagram, to determine the impact on the latter of motor design features and propellant characteristics. To achieve this goal, the method of analytical dependencies was developed. The equations obtained show that the required parameters of diagrams (including consumption-thrust characteristics difference between the starting and cruise modes) can be ensured due to varying either case diameter or propellant combustion rate or due to combined variation of these values. In practice, the cases are possible when for some reasons it does not seem possible to vary the case diameter or propellant combustion rate and the requirements to consumption diagram cannot be satisfied to the full extent. The task of motor developer in that case consists in determination of acceptable (alternative) propellant consumption diagrams that would be closest to required. The proposed method is based on calculation and construction of nomograms of dependencies of relative propellant consumption in cruse mode on relative time of starting leg at different propellant combustion rates and constant (required) case diameter and vice versa, at different values of case diameter and constant (available) propellant combustion rate. Using these nomograms, the rocket developer can determine the propellant consumption diagram acceptable for the rocket. In a number of cases, design limitations for separate main motor assemblies are imposed on consumption characteristic diagram that have an impact on its required parameters. The presented materials allow evaluating that impact and contain the proposals to remove it. The presented method allows quickly determining the conditions needed to fulfill required propellant combustion products consumption diagrams and in case of nonfulfillment of these conditions – allow presenting alternative options for selection of most acceptable one.

Key words: solid propellant charge mass, propellant combustion rate, combustion chamber pressure, operation time in starting and cruise modes, combustion chamber pressure difference

Bibliography:
1. K vyboru velichiny davliniia v kamere sgoraniia marshevykh RDTT: tekhn. otchet / GP “KB “Yuzhnoye”. Dnipro, 2017. 19 s.
2. Enotov V. G., Kushnir B. I., Pustovgarova Е. V. Avtomatizirovannaia proektnaia otsenka kharakteristik marshevykh dvigatelei na tverdom toplive s korpusom iz vysokoprochnykh metallicheskikh materialov takticheskikh i operativno-takticheskikh raket: ucheb.-metod. posobie / pod red. А. S. Kirichenko. Dnepropetrovsk, 2014. 72 s.
3. Sorkin R. Е. Gasotermodinamika raketnykh dvigatelei na tverdom toplive. М, 1967. 368 s.

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10.1.2020 Calculation and selection of parameters for a propellant consumption diagram of dual-thrust main SRM
10.1.2020 Calculation and selection of parameters for a propellant consumption diagram of dual-thrust main SRM
10.1.2020 Calculation and selection of parameters for a propellant consumption diagram of dual-thrust main SRM

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]]> 3.1.2020 Analysis of the unsteady stress-strain behavior of the launch vehicle hold-down bay at liftoff https://journal.yuzhnoye.com/content_2020_1-en/annot_3_1_2020-en/ Fri, 29 Sep 2023 18:22:49 +0000 https://journal.yuzhnoye.com/?page_id=32230
Inverse heat transfer method applied to capacitively cooled rocket thrust chambers. Thermal-structural analysis of regeneratively cooled thrust chamber wall in reusable LOX / Methane rocket engines.
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3. Analysis of the unsteady stress-strain behavior of the launch vehicle hold-down bay at liftoff

Authors:

Degtiarov М. O.1, Avramov K. V.2

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine1; Pidgorny A. Intsitute of Mechanical Engineering Problems, Kharkiv, Ukraine2

Page: Kosm. teh. Raket. vooruž. 2020, (1); 26-33

DOI: https://doi.org/10.33136/stma2020.01.026

Language: Russian

Annotation: The study of thermal strength of the hold-down bay is considered. The hold-down bay is a cylindrical shell with the load-bearing elements as the standing supports. The case of the hold-down bay consists of the following structural elements: four standing supports and the compound cylindrical shell with two frames along the top and bottom joints. The purpose of this study was the development of a general approach for the thermal strength calculation of the hold-down bay. This approach includes two parts. Firstly, the unsteady heat fields on the hold-down bay surface are calculated by means of the semi-empirical method, which is based on the simulated results of the combustion product flow of the main propulsion system. The calculation is provided by using Solid Works software. Then the unsteady stress-strain behavior of the hold-down bay is calculated, taking into consideration the plastoelastic deformations. The material strain bilinear diagram is used. The finiteelement method is applied to the stress-strain behavior calculation by using NASTRAN software. The thermal field is assumed to be constant throughout the shell thickness. As a result of the numerical simulation the following conclusions are made. The entire part of the hold-down bay, which is blown by rocket exhaust jet, is under stress-strain behavior. The stresses of the top frame and the shell are overridden the breaking strength that caused structural failure. The structure of the hold-down bay, which is considered in the paper, is unappropriated to be reusable. The hold-down bay should be reconstructed by reinforcement in order to provide its reusability.

Key words: stress-strain behavior, finite-element method, plastoelastic deformations, breaking strength, reusability

Bibliography:

1. Elhefny A., Liang G. Stress and deformation of rocket gas turbine disc under different loads using finite element modeling. Propulsion and Power Research. 2013. № 2. P. 38–49. https://doi.org/10.1016/j.jppr.2013.01.002
2. Perakis N., Haidn O. J. Inverse heat transfer method applied to capacitively cooled rocket thrust chambers. International Journal of Heat and Mass Transfer. 2019. № 131. P. 150–166. https://doi.org/10.1016/j.ijheatmasstransfer.2018.11.048
3. Yilmaz N., Vigil F., Height J., et. al. Rocket motor exhaust thermal environment characterization. Measurement. 2018. № 122. P. 312–319. https://doi.org/10.1016/j.measurement.2018.03.039
4. Jafari M. Thermal stress analysis of orthotropic plate containing a rectangular hole using complex variable method. European Journal of Mechanics A /Solids. 2019. № 73. P. 212–223. https://doi.org/10.1016/j.euromechsol.2018.08.001
5. Song J., Sun B. Thermal-structural analysis of regeneratively cooled thrust chamber wall in reusable LOX / Methane rocket engines. Chinese Journal of Aeronautics. 2017. № 30. P. 1043–1053.
6. Ramanjaneyulu V., Murthy V. B., Mohan R. C., Raju Ch. N. Analysis of composite rocket motor case using finite element method. Materials Today: Proceedings. 2018. № 5. P. 4920–4929.
7. Xu F., Abdelmoula R., Potier-Ferry M. On the buckling and post-buckling of core-shell cylinders under thermal loading. International Journal of Solids and Structures. 2017. № 126–127. P. 17–36.
8. Wang Z., Han Q., Nash D. H., et. al. Thermal buckling of cylindrical shell with temperature-dependent material properties: Conventional theoretical solution and new numerical method. Mechanics Research Communications. 2018. № 92. P. 74–80.
9. Duc N. D. Nonlinear thermal dynamic analysis of eccentrically stiffened S-FGM circular cylindrical shells surrounded on elastic foundations using the Reddy’s third-order shear de-formation shell theory. European Journal of Mechanics A /Solids. 2016. № 58. P. 10–30.
10. Trabelsi S., Frikha A., Zghal S., Dammak F. A modified FSDT-based four nodes finite shell element for thermal buckling analysis of functionally graded plates and cylindrical shells. Engineering Structures. 2019. № 178. P. 444–459.
11. Trinh M. C., Kim S. E. Nonlinear stability of moderately thick functionally graded sandwich shells with double curvature in thermal environment. Aerospace Science and Technology. 2019. № 84. P. 672–685.
12. Лойцянский Л. Г. Механика жидкости и газа. М., 2003. 840 с.
13. Launder B. E., Sharma B. I. Application of the energy dissipation model of turbulence to the calculation of flow near a spinning disc. International Journal of Heat and Mass Transfer. 1974. № 1. P. 131–138.
14. Михеев М. А., Михеева И. М. Основы теплопередачи. М., 1977. 345 с.
15. Малинин Н. Н. Прикладная теория пластичности и ползучести. М., 1968. 400 с.

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3.1.2020 Analysis of the unsteady stress-strain behavior of the launch vehicle hold-down bay at liftoff
3.1.2020 Analysis of the unsteady stress-strain behavior of the launch vehicle hold-down bay at liftoff
3.1.2020 Analysis of the unsteady stress-strain behavior of the launch vehicle hold-down bay at liftoff

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]]> 17.1.2020 Acoustic problems of rocket launch https://journal.yuzhnoye.com/content_2020_1-en/annot_17_1_2020-en/ Wed, 13 Sep 2023 11:36:44 +0000 https://journal.yuzhnoye.com/?page_id=31054
While saving the general thrust of the engine this method can not greatly reduce the sound radiation by a jet.
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17. Acoustic problems of rocket launch

Authors:

Hrinchenko V. T.

Organization:

Institute of Hydromechanics of National Academy of Sciences of Ukraine, Kyiv, Ukraine

Page: Kosm. teh. Raket. vooruž. 2020, (1); 155-159

DOI: https://doi.org/10.33136/stma2020.01.155

Language: Russian

Annotation: Due to an increase of power of rocket engines, the high intensity sound field generated by the exhaust jets have become an important factor, which determines the success rate of a rocket launch. Ensuring a successful launch of a rocket system became harder due to new engineering problems. Identification and definition of acoustic sources structure within a complex supersonic jet, being a one of the most important scientific problems, which have to be solved to find the ways to control accoustic radiation. A three components of acoustic sources can be defined here – broadband signals from large and small components of of turbulent jet and tonal signals which usually being overlooked during the estimation of overall sound pressure level. The paper considers various aspects of acoustics of the launch of rocket systems, which includes characteristics of acoustic sources in supersonic jets, possibilities and physical limitation factors, under which it is possible to control the sound radiation. Among the possible ways to control the process of sound generation by a jet, a method of water injection in a jet is being studied. While saving the general thrust of the engine this method can not greatly reduce the sound radiation by a jet. It is recommended to use big amounts of water-air mix to protect the launch pad from damage. Significant progress on the topic of understanding the process of sound generation by supersonic jets can be achieved via mathematical modeling of sound radiation. The latest achievements of mathematical modeling of sound generation by supersonic jets being presented.

Key words: Acoustics of rocket launch, acoustic efficiency of a jet, semi-empirical models of of jet acoustics, numeric-computational methods in aeroacoustics, control of jet-generated acoustic levels

Bibliography:
1. Lighthill M. J. On Sound Generated Aerodynamically: I. General Theory. Proc. Roy. Soc. London Ser. A, 211. 1952. Р. 564–581. https://doi.org/10.1098/rspa.1952.0060
2. Tam C. K. W. Jet noise. Theoretical Computftional Fluid Dynamics. 1998. No 10. Р. 393–405. https://doi.org/10.1007/s001620050072
3. Lubert C. P. Sixty years of launch vehicle acoustics. Proc.Mtgs.Acoust. Vol. 31. 2017. https://doi.org/10.1121/2.0000704
4. Ask the Astronaut: What does launch feel like? URL: https://www.airspacemag. com/ask-astronaut/ask-astronaut-what-does-launch-feel-what-thoughts-and-emotions-run-through-your-mind-180959920/
5. Tim P. Ask an Astronaut: My Guide to Life in Space. 2018. 272 p.
6. Saucer B. What’s the Deal with Rocket Vibration? MIT Technology Review. July 15, 2009. URL: https://www.technology-review.com/s/414364https:/whats-the-deal-with-rocket-vibrations/
7. Ross D. Mechanics of Underwater noise. 1976. 266 p.
8. Varnier J. Experimental study and simulation of rocket engine free jet noise. AIAA J. 2001. Vol. 39, Nо 10. P. 1851–1859. https://doi.org/10.2514/2.1199
9. Eldred K. M. Acoustic loads generated by the propulsion system. NASA SP-8072, 1971. 49 p.
10. Balakrishnan P., Srinivason K. Impinging get noise reduction using non-circular jets. Applied Acoustics. 2019. Vol. 143. Р. 19-30. https://doi.org/10.1016/j.apacoust.2018.08.016
11. Tsutsumi S. Acoustic generation mechanism of a supersonic jet impinging on deflectors / S. Tsutsumi, R. Takaki, Y. Nakanishi, K. Okamoto, S. Teramoto 52th AIAA Aerospace Sci. Meet. AIAA Pap. 2014-0882. 2014. 12 p. https://doi.org/10.2514/6.2014-0882
12. Ahuja K. K., Manes J. P., Massey K. C., Calloway A. B. An Evaluation of various concepts of Reducing Supersonic Jet Noise, AIAA-90-3982. AIAA 13th Aeroacoustic Conference, 1990. Р. 1-21. https://doi.org/10.2514/6.1990-3982
13. Krathapalli A., Lenkatakrishnan L., Elovarsan R., Laurenco L. Supersonic Jet Noise Suppression by Water Injection. AIAA 2000-2025. 6th AIAA/CEAS Aeroacoustic Conference, 2000. Р. 1-25.
14. Moratilla-Vega M. A., Lackhole K., Janicka J., Xia H., Page C. J. Jet Noise Analysis using an Efficient LES/ High-Order Acoustic Coupling Method. Computer and Fluid. 2020. https://doi.org/10.1016/j.compfluid.2020.104438

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17.1.2020 Acoustic problems of rocket launch
17.1.2020 Acoustic problems of rocket launch
17.1.2020 Acoustic problems of rocket launch

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]]> 12.2.2018 Methodological Support for Initial Phase Optimization of Projecting Design, Trajectory Parameters and Rocket Object Motion Control Programs https://journal.yuzhnoye.com/content_2018_2-en/annot_12_2_2018-en/ Thu, 07 Sep 2023 11:38:27 +0000 https://journal.yuzhnoye.com/?page_id=30770
The mathematical model of the guided missile provides adequate accuracy for design study to determine: overall dimensions and mass characteristics of the guided missile in general and its structural components and subsystems; power, thrust and consumption characteristics of the main engine; aerodynamic and ballistic characteristics of the guided missile. Flight Control Optimization and Thrust Optimization of Controllable Rocket Object Main Propulsion System.
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12. Methodological Support for Initial Phase Optimization of Projecting Design, Trajectory Parameters and Rocket Object Motion Control Programs

Authors:

Aksyonenko A. V.1, Baranov E. Y.1, Hurskyi O. I.1, Klochkov A. S.1, Morozov A. S.1, Alpatov A. P.2, Senkin V. S.2, Siutkina-Doronina S. V.2

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine1; The Institute of Technical Mechanics, Dnipro, Ukraine2

Page: Kosm. teh. Raket. vooruž. 2018 (2); 101-116

DOI: https://doi.org/10.33136/stma2018.02.101

Language: Russian

Annotation: The main scientific and methodological propositions for designing single-stage guided missiles with main solid rocket motors that are intended for delivering payload to the given spatial point with required and specified kinematic motion parameters are defined. The aim of the article is to develop methodology for the early design phase to improve the basic characteristics of guided missiles, including formalization of complex problem to optimize design parameters, trajectory parameters and motion control programs for guided missiles capable of flying along the ballistic, aeroballistic or combined trajectories. The task is defined as a problem of the optimal control theory with limitations in form of equality, inequality and differential constraints. An approach to program forming is proposed for motion control in the form of polynomial that brings the problem of the optimal control theory to a simpler problem of nonlinear mathematical programming. When trajectory parameters were calculated the missile was regarded as material point of variable mass and the combined equations for center-of-mass motion of the guided missile with projections on axes of the terrestrial reference system were used. The structure of the mathematical model was given along with the calculation sequence of criterion functional that was used for optimization of design parameters, control programs and basic characteristics of the guided missile. The mathematical model of the guided missile provides adequate accuracy for design study to determine: overall dimensions and mass characteristics of the guided missile in general and its structural components and subsystems; power, thrust and consumption characteristics of the main engine; aerodynamic and ballistic characteristics of the guided missile. The developed methodology was tested by solving design problems. Applications of the developed program were studied to present the research results in a user-friendly form.

Key words: complex problem of the optimal control theory, problem of nonlinear mathematical programming, main solid rocket motor, limitations for motion parameters and basic characteristics of the object

Bibliography:
1. Degtyarev A. V. Rocket Engineering: Problems and Prospects. Selected scientific-technical publications. Dnepropetrovsk, 2014. 420 p.
2. Shcheverov D. N. Designing of Unmanned Aerial Vehicles. М., 1978. 264 p.
3. Sinyukov А. М. et al. Ballistic Solid-Propellant Rocket / Under the editorship of A. M. Sinyukov. М., 1972. 511 p.
4. Varfolomeyev V. I. Designing and Testing of Ballistic Rockets / Under the editorship of V. I. Varfolomeyev, M. I. Kopytov. М., 1970. 392 p.
5. Vinogradov V. A., Grushchansky V. A., Dovgodush S. I. et al. Effectiveness of Complex Systems. Dynamic Models. М., 1989. 285 p.
6. Il’ichyov A. V., Volkov V. D., Grushchansky V. A. Effectiveness of Designed Complex Systems’ Elements. М., 1982. 280 p.
7. Krotov V. F., Gurman V. I. Methods and Problems of Optimal Control. М., 1973. 446 p.
8. Pontryagin L. S. et al. Mathematical Theory of Optimal Processes. М., 1969. 385 p.
9. Tarasov E. V. Algorithms of Flying Vehicles Optimal Designing. М., 1970. 364 p.
10. Alpatov A. P., Sen’kin V. S. Complex Task of Optimization of Space Rocket Basic Design Parameters and Motion Control Programs. Technical Mechanics. 2011. No. 4. P. 98-113.
11. Alpatov A. P., Sen’kin V. S. Methodological Support for Selection of Launch Vehicle Configuration, Optimization of Design Parameters and Flight Control Programs. Technical Mechanics. 2013. No. 4. P. 146-161.
12. Sen’kin V. S. Optimization of Super-Light Launch Vehicle Design Parameters. Technical Mechanics. 2009. No. 1. P. 80-88.
13. Sen’kin V. S. Flight Control Optimization and Thrust Optimization of Controllable Rocket Object Main Propulsion System. Technical Mechanics. 2000. No. 1. P. 46-50.
14. Syutkina-Doronina S. V. On Problem of Optimization of Design Parameters and Control programs of a Rocket Object With Solid Rocket Motor. Aerospace Engineering and Technology. 2017. No. 2 (137). P. 44-59.
15. Lebedev А. А., Gerasyuta N. F. Rocket Ballistics. М., 1970. 244 p.
16. Razumov V. F., Kovalyov B. K. Design Basis of Solid-Propellant Ballistic Missiles. М., 1976. 356 p.
17. Yerokhin B. T. SRM Theoretical Design Basis. М., 1982. 206 p.
18. Abugov D. I., Bobylyov V. M. Theory and Calculation of Solid Rocket Motors. М., 1987. 272 p.
19. Shishkov А. А. Gas Dynamics of Powder Rocket Motors. М., 1974. 156 p.
20. Sen’kin V. S. Complex Task of Optimization of Super-Light Solid-Propellant Launch Vehicle Design Parameters and Control Programs. Technical Mechanics. 2012. No. 2. P. 106-121.
21. Methodological Support to Determine in Initial Designing Phase the Design Parameters, Control Programs, Ballistic, Power, and Mass-Dimensional Characteristics of Controllable Rocket Objects Moving In Aeroballistic Trajectory: R&D Report. ITM of NASU and SSAU, Yuzhnoye SDO. Inv. No. 40-09/2017. 2017. 159 p.

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12.2.2018 Methodological Support for Initial Phase Optimization of Projecting Design, Trajectory Parameters and Rocket Object Motion Control Programs
12.2.2018 Methodological Support for Initial Phase Optimization of Projecting Design, Trajectory Parameters and Rocket Object Motion Control Programs
12.2.2018 Methodological Support for Initial Phase Optimization of Projecting Design, Trajectory Parameters and Rocket Object Motion Control Programs

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]]> 10.2.2018 Calculation of Gas Flow in High-Altitude Engine Nozzle and Experience of Using Water-Cooled Nozzle Head during Tests https://journal.yuzhnoye.com/content_2018_2-en/annot_10_2_2018-en/ Thu, 07 Sep 2023 11:29:45 +0000 https://journal.yuzhnoye.com/?page_id=30766
Thrust-Vectoring Nozzle Performance Mode-ling.
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10. Calculation of Gas Flow in High-Altitude Engine Nozzle and Experience of Using Water-Cooled Nozzle Head during Tests

Authors:

Nikitenko K. O.

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine

Page: Kosm. teh. Raket. vooruž. 2018 (2); 83-93

DOI: https://doi.org/10.33136/stma2018.02.083

Language: Russian

Annotation: At Yuzhnoye State Design Office, the Cyclone-4 launch vehicle 3rd stage engine has been developed and is under testing. For adjustment of the engine and test bench systems, in the first firing tests the radiation-cooled nozzle extension was replaced with a steel water-cooled one. It was planned to start the engine with water-cooled nozzle extension without vacuumizing and without gad dynamic pipe, which conditioned operation with flow separation at the output edge of water-cooled nozzle extension. Therefore, the calculation of flow in the nozzle with water-cooled extension, flow separation place, and thermal load on watercooled nozzle extension during operation in ground conditions is an important task. Selection of turbulent flow model has a noticeable impact on prediction of flow characteristics. The gas dynamic analysis of the nozzle with water-cooled extension showed the importance of using the turbulent flow model k-ω SST for the flows with internal separation of boundary layer and with flow separation at nozzle section. The use the flow model k-ω SST for calculation of nozzle with flow separation or with internal transitional layer allows adequately describing the flow pattern, though, as the comparison with experimental data showed, this model predicts later flow separation from the wall than that obtained in the tests. The calculation allows obtaining a temperature profile of the wall and providing the recommendations for selection of pressure measurement place in the nozzle extension for the purpose of reducing sensors indication error. With consideration for the special nature of the nozzle extension wall temperature field, the cooling mode was selected. The tests of RD861K engine nozzle with water-cooled extension allow speaking about its successful use as a required element for testing engine start and operation in ground conditions without additional test bench equipment.

Key words: turbulent flow, flow separation, cooling, technological extension

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2. Mezhevov A. V., Skoromnov V. I., Kozlov A. V. et al. Introduction of Radiation Cooling Nozzle Head of Made of Carbon-Carbon Composite Material on DM-SL Upper Stage 11D58M Main Engine. News of Samara Aerospace University. No. 2 (10). 2006. P. 260-264.
3. Fluent. Software Package, Ver. 6.2.16, Fluent Inc., Lebanon, NH, 2004.
4. Wilcox D. C. Turbulence Modeling for CFD. DCW Industries, Inc. La Canada, California, 1998. 460 р.
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7. Rajasekaran A., Babu V. Numerical Simulation of Three-dimensional Reacting Flow in a Model Supersonic Combustor. Journal of Propulsion and Power. Vol. 22. No. 4. 2006. Р. 820-827. https://doi.org/10.2514/1.14952
8. Spalart P., Allmaras S. A one-equation turbulence model for aerodynamic flows: Technical Report. American Institute of Aero-nautics and Astronautics. AIAA-92-0439. 1992. Р. 5-21. https://doi.org/10.2514/6.1992-439
9. Launder B. E., Spalding D. B. Lectures in Mathematical Models of Turbulence. London, 1972. Р. 157-162.
10. Rajasekaran A., Babu V. Numerical Simulation of Three-dimensional Reacting Flow in a Model Supersonic Combustor. Journal of Propulsion and Power. Vol. 22. No. 4. 2006. Р. 820-827. https://doi.org/10.2514/1.14952
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10.2.2018 Calculation of Gas Flow in High-Altitude Engine Nozzle and Experience of Using Water-Cooled Nozzle Head during Tests
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]]> 5.2.2018 Electromagnetic Valves Developed by Yuzhnoye SDO Liquid Rocket Engines Design Office https://journal.yuzhnoye.com/content_2018_2-en/annot_5_2_2018-en/ Thu, 07 Sep 2023 11:01:49 +0000 https://journal.yuzhnoye.com/?page_id=30749
The present-day spacecraft gas-jet orientation and stabilization systems use as propulsion devices the electromagnetic valves with nozzles whose thrust is, as a rule, not more than 30 N and the working medium pressure is up to 24 MPa. Yuzhnoye State Design Office developed for 15B36 gas-jet system the electropneumatic valve with amplification and nozzle, which is operable at the pressure below 45 MPa, ensures the action frequency of up to 10 Hz and is capable of creating the thrust of 100 N on gaseous argon. Recently, Yuzhnoye State Design Office develops the apogee RD840 LRE with 400 N thrust, for the conditions of which the direct-action electrohydraulic valve was developed and tested with the following characteristics: pressure – up to 2.15 MPa, consumed power in operation mode – less than 7.1 W, action time – not more than 0.02 s, mass – 0.19 kg.
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5. Electromagnetic Valves Developed by Yuzhnoye SDO Liquid Rocket Engines Design Office

Authors:

Konokh V. I., Boiko V. S., Troyak A. B., Ivashura A. V.

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine

Page: Kosm. teh. Raket. vooruž. 2018 (2); 34-48

DOI: https://doi.org/10.33136/stma2018.02.034

Language: Russian

Annotation: In the pneumohydraulic systems of liquid rocket engines and propulsion systems, electromagnetic valves that allow making the pneumohydraulic systems more simple and ensuring multiple ignition of liquid rocket engines have found wide application. The Yuzhnoye-developed electromagnetic valves are designed according to two schemes – of direct and indirect action. In the direct-action electromagnetic valves, the shutting-off device opens (closes) the throat with the force developed by electric magnet. They have gained acceptance in the pneumohydraulic systems with the working medium pressure of ~8.5 MPa, they are of simple design and have high operating speed (0.001…0.05 s). In the electromagnetic valves with amplification, the electromagnet armature is connected with control valve and the main shutting-off device moves due to the force from working medium pressure drop on it. They are used in the operating pressure range of 0.5…56 MPa, at that, the action time is 0.025…0.15 s. For the European Vega launch vehicle fourth stage main engine assembly that has pressure propellant feeding system, the electrohydraulic valve with amplification and drainage was developed. The dependence of this electrohydraulic valve high speed from the line’s output length is decreased to the maximum due to the installation of Venturi nozzle at the output connecting branch. This electrohydraulic valve is operable at the pressure below 8 MPa, the action time is 0.08…0.12 s. The present-day spacecraft gas-jet orientation and stabilization systems use as propulsion devices the electromagnetic valves with nozzles whose thrust is, as a rule, not more than 30 N and the working medium pressure is up to 24 MPa. Yuzhnoye State Design Office developed for 15B36 gas-jet system the electropneumatic valve with amplification and nozzle, which is operable at the pressure below 45 MPa, ensures the action frequency of up to 10 Hz and is capable of creating the thrust of 100 N on gaseous argon. To solve the task of decreasing the dependence of operability and high speed of electromagnetic valves with drainage and amplification on geometry of lines in which a valve is installed, the electropneumatic valve was developed that has spool elements ensuring reliable and quick action with long input lines of 0.004 m diameter. Its mass is 2…2.5 times lower than the mass of analogs. Recently, Yuzhnoye State Design Office develops the apogee RD840 LRE with 400 N thrust, for the conditions of which the direct-action electrohydraulic valve was developed and tested with the following characteristics: pressure – up to 2.15 MPa, consumed power in operation mode – less than 7.1 W, action time – not more than 0.02 s, mass – 0.19 kg. The presented electromagnetic valves by their technical and operational characteristics meet the highest world requirements and have found wide utility in liquid rocket engines and propulsion systems.

Key words: electrohydraulic valve, electropneumatic valve, pneumohydraulic system, direct-action electric valve, electric valve with amplification, action time

Bibliography:
1. Electric Hydraulic Valve: Patent 89948 Ukraine: MPK F 16K 32/02 / Shnyakin V. M., Konokh V. I., Kotrekhov B. I., Troyak A. B., Boiko V. S.; Applicant and patent holder Yuzhnoye State Design Office. а 2006 02543; claimed 09.03.2006; published 25.03.2010, Bulletin No. 6.
2. Boiko V. S., Konokh V. I. Stabilization of Opening Time of Electric Hydraulic Valve with Boost in Liquid Rocket Engine Hydraulic System. Problems of Designing and Manufacturing Flying Vehicle Structures: Collection of scientific works. 2015. Issue 4 (84). P. 39-48.
3. Electric Valve: Patent 97841, Ukraine: MPK F 16K 32/02 / Shnyakin V. M., Konokh V. I., Kotrekhov B. I., Troyak A. B., Boiko V. S., Ivashura A. V.; Applicant and patent holder Yuzhnoye State Design Office. а 2009 12002; claimed 23.11.2009; published 26.03.2012, Bulletin No. 6.
4. Boiko V. S., Konokh V. I. Increase of Action Stability of Electric Pneumatic Valve with Boost in the System with Increased Inlet Hydraulic Resistance. Aerospace Engineering and Technology: Scientific-Technical Journal. 2013. Issue 3 (100). P. 90-95.
5. Flying Vehicles Pneumatic Systems Units / Lyaskovsky I. F., Shishkov A. I., Romanenko N. T., Romanenko M. T., Chernov M. T., Yemel’yanov V. V. / Under the editorship of N. T. Romanenko. М., 1976. 176 p.

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5.2.2018 Electromagnetic Valves Developed by Yuzhnoye SDO Liquid Rocket Engines Design Office
5.2.2018 Electromagnetic Valves Developed by Yuzhnoye SDO Liquid Rocket Engines Design Office
5.2.2018 Electromagnetic Valves Developed by Yuzhnoye SDO Liquid Rocket Engines Design Office

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