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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

Вацлав Самусевич

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

Bibliography:

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22. Aksenenko A. V., Baranov E. Yu., Hursky A. I., Klochkov A. S., Morozov A. S., Alpatov A. P., Senkin V. S., Siutkina-Doronina S. V. Metodicheskoe obespechenie dlia optimizatsii na nachalnom etape proektirovaniia proektnykh parametrov, parametrov traektorii i programm upravleniia dvizheniem raketnogo ob’ekta. Kosmicheskaia tekhnika. Raketnoe vooruzhenie: sb. nauch.-tekhn. st. / GP “KB “Yuzhnoye”. Dnipro, 2018. Vyp. 2 (116). S. 101–116. https://doi.org/10.33136/stma2018.02.101

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12.1.2024 Hardening of steels modifying their surfaces with ion-plasma nitriding in glow discharge

manager — Mon, 17 Jun 2024 11:36:02 +0000

12. Hardening of steels modifying their surfaces with ion-plasma nitriding in glow discharge

Authors:

Nadtoka V. M.¹, Shtapenko Ye. P.², Kraeva V. S.¹, Kraev M. V.¹

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine¹; Ukrainian State University of Science and Technologies²

Page: Kosm. teh. Raket. vooruž. 2024, (1); 102-113

DOI: https://doi.org/10.33136/stma2024.01.102

Language: Ukrainian

Annotation: Steel hardening technology is considered, which implies modification of the steel surface with the method of ion-plasma nitriding in glow discharge. Ion-plasma nitriding is a multi-factor process, which requires the study of the influence of nitriding process conditions on the structure of modified layers, which, in its turn, determines their mechanical properties. The subjects of research included: austenitic steel 12X18Н10T, carbon steel Ст3 and structural steel 45. There were two conditions of plasma creation during the research: free location of samples on the surface of the cathode (configuration I) and inside the hollow cathode (configuration II). Optimal parameters of the ion-plasma nitriding process have been determined, which provide stability of the process and create conditions for intensive diffusion of nitrogen into the steel surface. Hydrogen was added to the argon-nitrogen gaseous medium to intensify the nitriding process. Working pressure in the chamber was maintained within the range of 250-300 Pa, the duration of the process was 120 minutes. Comparative characteristics of the structure and microhardness of the modified surfaces of the steels under study for two ion-plasma nitriding technologies are presented. Metallographic examination of the structure of the surface modified layers in the cross section showed the presence of the laminated nitrided layer, which consists of different phases and has different depths, depending on the material of the sample and treatment mode. Nitrided layer of 12Х18Н10Т steel consisted of four sublayers: upper “white” nitride layer, double diffuse layer and lower transition layer. The total depth of the nitrided layer after the specified treatment time reached 23 μm, use of hollow cathode increased it by 26% to 29 μm. The nitrided layers of steel Ст3 and steel 45 consisted of two sublayers – thick “white” nitride layer and general diffuse layer with a thickness of about 18 μm. The microhardness of the nitrided layer of steel Ст3 was 480 HV, increasing by 2,5 times, and for steel 45 was 440 HV, increasing by 1,7 times. The use of hollow cathode for these steels reduces the depth of the nitrided layer, but at the same time the microhardness increases due to the formation of a thicker and denser nitride layer on the surface. The results of the conducted research can be used to strengthen the surfaces of the steel parts in rocket and space technology, applying high-strength coatings.

Key words: ion nitriding, glow discharge, cross-sectional layer structure, hardening, microhardness

Bibliography:

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

Дмитрий Макренко — Mon, 17 Jun 2024 08:41:58 +0000

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

Bibliography:

Arkhipov V. Erosionnoe gorenie condensirovannykh system. Sb. tr. ІХ Vserossiyskoy nauch. conf. 2016 g. (FPPSM-2016). Tomsk, 2016.
Mukunda S., Paul P. J. Universal behaviour in erosive burning of solid propellants. Combustion and flame, 1997. https://doi.org/10.1016/S0010-2180(96)00150-2
Sabdenov K. , Erzda M., Zarko V. Ye. Priroda i raschet skorosti erozionnogo goreniya tverdogo raketnogo topliva. Inzhenerniy journal: nauka i innovatsii, 2013. Vyp. 4.
Evlanova A., Evlanov A. A., Nikolaeva Ye. V. Identifikatsiya parametrov erozionnogo goreniya topliva po dannym ognevykh stendovykh ispytaniy. Izvestiya TulGU. Tekhn. nauki. 2014. Vyp. 12, ch. 1.
Yanjie Ma, Futing Bao, Lin Sun, Yang Liu, and Weihua Hui. A New Erosive Burning Model of Solid Propellant Based on Heat Transfer Equilibrium at Propellant Surface. Hindawi International Journal of Aerospace Engineering, Vol. 2020, Article ID 8889333. https://doi.org/10.1155/2020/8889333
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Stepanov Ye. A. Gasodynamicheskie funktsii. Mashinostroenie, Moskva, 1965.
William Orvis. EXCEL dlya uchenykh, inzhenerov i studentov. Kiev: «Junior», 1999.

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15.1.2024 Enhancing operability of the fuel system units in the hot climate conditions

manager — Mon, 17 Jun 2024 07:43:36 +0000

15. Enhancing operability of the fuel system units in the hot climate conditions

Authors:

Udod A. M.¹, Skokov O. I.¹, Volovschikova V. V.¹, Shepel S. V.², Litvinova M. V.², Steshenko K. A.¹

Organization:

DINTEM Ukrainian Research Design-Technological Institute of Elastomer Materials and Products LLC¹; FED Joint Stock Company²

Page: Kosm. teh. Raket. vooruž. 2024, (1); 129-135

DOI: https://doi.org/10.33136/stma2024.01.129

Language: Ukrainian

Annotation: The article dwells on the problem of enhancement of durability for the mechanical rubber articles, which is directly related to the enhance of rubber resistance to various types of heat aging. Heat resistance during compression is most important for rubbers used for seals of various types: rings, collars, armored collars, gaskets for aviation and rocket technology hardware. Stress relaxation and the accumulation of relative residual deformation of rubbers, caused by the kinetic rearrangement of chemical bonds, are extremely sensitive to the influence of high temperatures. The main cause of the defects is the loss of elastic properties of the seals because of the accelerated heat aging of the nitrile group under conditions of long-term exposure to elevated temperatures in conditions of hot climate. The results of accelerated climatic testing of specimens of mechanical rubber articles, as well as the results of climatic endurance testing of the units for the period simulating 20-year service life are specified, and the main types of defects which result in the loss of performance properties of the mechanical rubber articles are as follows: great (up to 100%) residual deformation of intersections, cracking, loss of elasticity. The warranty life of fuel system units, made of ИРП-1078 nitrile rubber, does not exceed 12 years. Replacing the existing rubbers with rubbers created on the basis of more heat-bearing rubbers is the most promising way to improve the performance properties of the mechanical rubber articles under the high temperatures. The new D2301 rubber is based on fluorosiloxane rubber. It provides high thermal stability and, especially, the ability to maintain high performance properties for a long time under the simultaneous impact of hostile environment and high temperatures. The results of climatic endurance testing of fuel system units, equipped with rubber articles made of D2301 rubber, fully justify the increase of the specified service life of the specified units from 12 to 16 years. It is recommended to introduce D2301 rubber into the effective normative documentation and continue studies in order to extend the nomenclature of mechanical rubber articles made of D2301 rubber to provide the reliable sealing of units during the service life of 16 years or longer.

Key words: leaktightness of articles, fluorosiloxane rubber, rubber, temperature of the hot climate, physical-mechanical properties of the rubber, climatic endurance tests, elastic properties, warranty life

Bibliography:

Lepetov V. A., Yurtsev L. N. Raschet i konstruirovanie rezinovykh izdeliy. Moskva.
Khimia. 1971. 417 s.

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

manager — Fri, 14 Jun 2024 11:36:31 +0000

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.

Bibliography:

Pankratov Yu. , Novikov A. V., Tatarevsky K. E., Azanov I. B. Dynamika perekhodnykh processov. 2014.
Sinyukov A. M., Morozov N. I. Konstruktsia upravlyaemykh ballisticheskykh raket. 1969.
Kabakova Zh. V., Kuda S. A., Logvinenko A. I., Khomyak V. A. Opyt razrabotki pneumosystemy dlya otdelenita golovnogo aerodynamicheskogo obtekatelya. Kosmicheskaya technika. Raketnoe vooruzhenie. 2017. Vyp. 2 (114).
Kolesnikov K. S., Kozlov V. V., Kokushkin V. V. Dynamika razdeleniya stupeney letatelnykh apparatov. 1977.
Antares – Spaceflight Insider: web site. URL: https://www. Spaceflightinsider.com/missions/iss/ng-18-cygnus-cargo-ship-to-launch-new-science-to-iss/Antares (data zvernennya 30.10.2023).
Falcon 9 – pexels: website. URL: https://www. pexels.com/Falcon 9 (data zvernennya 31.10.2023).
Kolesnikov K. , Kokushkin V. V., Borzykh S. V., Pankova N. V. Raschet i proektirovanie system razdeleniya stupeney raket. 2006.
Cyclone-4M – website URL: https://www.yuzhnote.com (data zvernennya 31.10.2023)
Logvinenko A. Sozdanie gasoreaktivnykh system otdeleniya i uvoda otrabotavshykh stupeney – noviy shag v RKT. Kosmicheskaya tekhnika. Raketnoe vooruzhenie, KBU, NKAU, vyp. 1, 2001.
Logvinenko A. I., Porubaimekh V. I., Duplischeva O. M. Sovremennye metody ispytaniy system i elementov konstruktsiy letatelnykh apparatov. Monografia. Dnepr, KBU, 2018.

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2.1.2024 New and advanced liquid rocket engines of the Yuzhnoye SDO

Дмитрий Макренко — Wed, 12 Jun 2024 15:04:41 +0000

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:

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

Вацлав Самусевич

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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10.2.2019 Dynamic performance of the gas drive with jet motor

Вацлав Самусевич — Tue, 03 Oct 2023 11:52:15 +0000

10. Dynamic performance of the gas drive with jet motor

Authors:

Oliinyk V. P., Kaluger L. G.

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine

Page: Kosm. teh. Raket. vooruž. 2019, (2); 71-79

DOI: https://doi.org/10.33136/stma2019.02.071

Language: Russian

Annotation: The use of servo drives on flying vehicles determines the requirements to their dynamic characteristics. The problems of dynamics of drive with jet motor are not practically covered in publications. The task arises of selection of structure and parameters of devices consisting of several subsystems whose dynamic characteristics must be brought into agreement with each other in optimal way. The purpose of this work is to develop mathematical dependences for calculation of dynamic characteristics. The functional arrangement of the drive is considered consisting of jet motor based on Segner wheel with de Laval nozzle, mechanical transmission, pneumatic distributing device – jet pipe controlled by electromechanical converter. The layout is presented of mechanical segment of servo drive with jet motor with screw-nut transmission. The dynamic model is presented and the algebraic relations to determine natural frequencies of the drive are given. The motion equations of output rod at full composition of load are given. Using Lagrange transformation as applied to ball screw transmission, the expression for reduced mass of output element was derived. The reduced mass of load depends on the jet motor design and exerts basic influence on the drive’s natural frequencies. The evaluation is given of reduced mass change from the jet motor moment of inertia and reducer transmission coefficient. Based on the proposed algorithms, the dynamic characteristics of servo drive were constructed: transient process and amplitude-frequency characteristic. The drive has relatively low pass band, which is explained by the value of reduced mass of load.

Key words: pneumatic drive, functional arrangement, hydrodynamic force, reduced mass, Lagrange transformations, ball screw transmission, transient process, frequency characteristic

Bibliography:

1. Pnevmoprivod system upravleniya letatelnykh apparatov /V. A. Chaschin, O. T. Kamladze, A. B. Kondratiev at al. M., 1987. 248 s.

2. Berezhnoy A. S. Sovershenstvovanie rabochikh characteristic struino-reaktivnogo pnevmoagregata na osnove utochneniya modeli rabochego processa: dis. cand. techn. nauk: 05.05.17. Zaschischena 03.10.14. Sumy, 2014. 157 s.

3. Oleinik V. P., Yelanskiy Yu. A., Kovalenko V. N. et al. Staticheskie characteristiki gazovogo privoda so struinym dvigatelem /Kosmicheskaya technika. Raketnoe vooruzhenie: Sb. nauch.-techn. st. 2016. Vyp. 2. S. 21-27.

4. Abramovich G. N. Prikladnaya gazovaya dynamika. M., 1976. 888 s.

5. Strutinskiy V. B. Matematichne modelyuvannya processiv ta system mechaniki. Zhitomir, 2001. 612 s.

6. Shalamov A. V., Mazein P. G. Dynamicheskaya model’ sharikovintovoi pary/ Izv. Chelyabinskogo nauchnogo centra UrO RAN. №4. Chelyabinsk, 2002. S.161-170.

7. Kripa K.Varanasi, Samir A. Nayfer. The Dynamics of Lead-Screw Drivers: Low-Order Modeling and Experiments /Journal of Dynamic System, Measurement and Control. June 2004. Vol. 126. P. 388-395. https://doi.org/10.1115/1.1771690

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9.2.2019 Gas-dynamic simulation of the supersonic stream in the pulsed wind tunnel

manager — Tue, 03 Oct 2023 11:48:27 +0000

9. Gas-dynamic simulation of the supersonic stream in the pulsed wind tunnel

Authors:

Sirenko V. M., Zhivotov O. Yu.

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine

Page: Kosm. teh. Raket. vooruž. 2019, (2); 63-70

DOI: https://doi.org/10.33136/stma2019.02.063

Language: Russian

Annotation: A promising experimental bench – a shock wind tunnel was put into operation at Yuzhnoye SDO. The shock wind tunnel is designed to simulate the incident flow during rocket flight at high supersonic and hypersonic velocities. To solve actual design problems facing Yuzhnoye SDO, it was necessary to expand the range of velocities under investigation in a shock wind tunnel by low supersonic Mach numbers (Mа=1.5; 2; 3). As a result of this work, a modernized configuration of the shock wind tunnel was developed, which allows simulating flow parameters at low supersonic velocities. The results of aerodynamic experiment performed in the modernized shock wind tunnel, which are close to full scale ones, can be obtained using as much data on peculiarities of supersonic flow formation in it as possible. Therefore, the study of the distribution of Mach numbers profiles in the working section of the modernized shock wind tunnel at low and high supersonic velocity was chosen as the main line of research. The results of the research presented in the article are based on the use of numerical simulation methods, as well as data obtained experimentally. As a result of gasdynamic simulation of a supersonic flow conducted for the nozzle Mа=4 and the nozzle Mа=2, the calculated and experimental data on the distribution pattern and field values of Mach numbers in the working section of the tunnel were obtained. A comparative analysis was carried out. The boundaries of the region of equal velocities, within which the condition of quasistatic supersonic flow is satisfied, and the lifetime of the operating mode for the selected nozzle type were determined. At the flow from the nozzle Mа=2, a peculiarity was revealed in the distribution pattern of Mach numbers fields associated with the appearance of “blocking” effect of the supersonic flow. The methods for eliminating the effect of flow “blocking” at low supersonic velocities are proposed.

Key words: incident flow modeling, velocity fields in the wind tunnel working section, aerodynamic experiment

Bibliography:

1. Zvegintsev V. I. Gasodynamicheskie ustanovki kratkovremennogo deistviya. V dvuh chastyakh. Ch. 1. Ustanovki dlya nauchnykh issledovaniy. Novosibirsk, 2014. 551 s.

2. Computerno-vymiryuvalni tekhnologii kontrolu ta upravlinnya raketno-kosmichnoi techniki / monogr. pid zagal. red. prof. V. P. Malaichuka. Dnipro, 2018. 344 s.

3. «Sirius-18». Systema izmereniya i upravleniya impulsnoi aerodynamicheskoi truboi. Rukovodstvo po ekspluatatsii. ELVA4.044.901 RE. 2018. 45 s.

4. Abramovich G.N. Prikladnaya gazovaya dynamika. M., 1978. 888 s.

5. Raschet vnutrennego davlenia v otsekakh RN. YSF YZH UMN 041 01. Rukovodstvo operatora. 2016. 138 s.

6. Issledovania characteristic hyperzvukovoi aerodynamicheskoi truby AT-303. Ch. 1. Polya skorostey / A. M. Kharitonov at al. Teplophysika i aeromekhanika. 2006. T. 13, № 1. S. 1–17.

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

manager — Fri, 29 Sep 2023 18:22:49 +0000

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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USA	Boardman; Matawan; Boydton; Plano; Miami; Columbus; Columbus; Columbus; Detroit; Phoenix; Phoenix; Phoenix; Monroe; Ashburn; Seattle; Ashburn; Boardman; Seattle; Portland; San Mateo; Des Moines; Boardman; Boardman; Ashburn; Ashburn	25
Singapore	Singapore; Singapore; Singapore; Singapore; Singapore; Singapore; Singapore; Singapore; Singapore; Singapore; Singapore	11
Ukraine	Dnipro; Kyiv; Dnipro	3
Germany	;; Falkenstein	3
Finland	Helsinki	1
Great Britain	London	1
Canada	Monreale	1
Romania	Voluntari	1
Netherlands	Amsterdam	1

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