Search Results for “launch site” – Collected book of scientific-technical articles

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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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:

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5.1.2024 Assessment of risk of toxic damage to people in case of a launch vehicle accident at flight

Дмитрий Макренко — Thu, 13 Jun 2024 06:00:42 +0000

5. Assessment of risk of toxic damage to people in case of a launch vehicle accident at flight

Authors:

Hladkyi E. H., Sheiko А. F.

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine

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)].
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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:

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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.
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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.
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15. Малинин Н. Н. Прикладная теория пластичности и ползучести. М., 1968. 400 с.

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9.1.2020 Experimental investigation of a liner-free propellant tank made from polymer composite materials

Вацлав Самусевич — Wed, 13 Sep 2023 10:43:08 +0000

9. Experimental investigation of a liner-free propellant tank made from polymer composite materials

Authors:

Sidoruk А. V., Popov D. А., Zadoia А. S., Kalinichenko D. S., Aksenenko O. V., Husarova I. O., Derevianko I. I., Kharchenko V. M., Litot O. V.

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine

Page: Kosm. teh. Raket. vooruž. 2020, (1); 90-98

DOI: https://doi.org/10.33136/stma2020.01.090

Language: Russian

Annotation: The exploratory and experimental investigations were conducted into design of propellant tank made of composite polymer materials for work in cryogenic environment at operating pressure of 7.5 kgf/cm2 . When determining the configuration of a liner-free composite propellant tank, the main requirement was ensuring its leak-tightness at internal excess pressure and cryogenic temperature effect. The world experience of creating similar designs was analyzed and the requirements were defined imposed on configuration of propellant tank load-bearing shells. Before defining the final configuration, the types of materials, reinforcing patterns, and possible ways to ensure leak-tightness were analyzed, and preliminary tests were conducted of physical and mechanical characteristics of thin-wall samples of composite materials and tubular structures with different reinforcing patterns. The tests of carbon plastic samples were conducted at different curing modes to determine the most effective one from the viewpoint of strength characteristics and the tests for permeability by method of mouthpiece were conducted. The tests of pilot propellant tank showed that the calculated values of deformations and displacements differ from the experimental values by no more than 10 %. Using the parameters measurement results from the tests on liquid nitrogen, the empirical formulas were obtained to calculate linear thermal expansion coefficient of the package of materials of load -bearing shell. The empirical dependences were constructed of relative ring deformations at load-bearing shell middle section on pressure and temperature. The tests confirmed correctness of adopted solutions to ensure strength and leak-tightness of propellant tank load-bearing shell at combined effect on internal excess pressure and cryogenic temperature, particularly at cyclic loading. The materials used and propellant tank manufacturing technologies ensure leak-tightness of load-bearing shell at liquid nitrogen operating pressure of 7.5 kgf/cm2 and strength at excess pressure of 15 kgf/cm2 and allow conducting approbation of prospective stage of the integrated launch vehicle.

Key words: load-bearing shell, permeability, cryogenic propellant, relative deformations, linear thermal expansion coefficient

Bibliography:

1. Frantsevich I. М., Karpinos D. М. Kompozitsionnye materialy voloknistogo stroeniia. K., 1970.

2. TSM YZH ANL 009 00. Composite fuel tank for ILV, Dnipro, Yuzhnoye SDO, 2019.

3. Zheng H., Zeng X., Zhang J., Sun H. The application of carbon fiber composites in cryotank. Solidification. 2018. https://doi.org/10.5772/intechopen.73127

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7.1.2020 Studying the motion of a launch vehicle and observed space debris objects during launch preparation

Вацлав Самусевич — Wed, 13 Sep 2023 06:27:07 +0000

7. Mechanics of a satellite cluster. Methods for estimating the probability of their maximal approach in flight

Authors:

Holubek O. V.

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine

Page: Kosm. teh. Raket. vooruž. 2020, (1); 76-84

DOI: https://doi.org/10.33136/stma2020.01.076

Language: Russian

Annotation: The mathematic modeling was performed of the flight of light-class three-stage launch vehicle injecting a payload into sun-synchronous orbit of 700 km altitude and a cluster of observed space debris objects in the conditions of dynamically changing cataloged space situation. It is shown that as the launch moment becomes closer, the cataloged space situation is ascertained, which leads to the constant change of the quantity of hazardous space debris objects observed in the vicinity of launch vehicle trajectory and to the change of the parameters of their approach to the launch vehicle: minimal relative distance, relative velocity, rendezvous angle and launch moment for which hazardous approach is revealed. The hazardous approaches for the launch vehicle trajectory under consideration are more often observed with the relative velocities of more than 8 km/s and rendezvous angles less than 90 deg and their variations within the launch window do not exceed 1.2 m/s and 0.035 deg respectively. In this case, the histograms of distribution of relative distance, relative velocity, and rendezvous angle from catalog to catalog vary insignificantly. The distribution of hazardous approaches in launch time within launch window is not uniform, the regions are observed with low quantity of hazardous approaches and with high quantity. The hazard of launch vehicle collision with observed space debris objects in a launch is confirmed. In all, in the launch day time window under consideration, more than ten hazardous approaches are revealed, for two of them the approach to minimal distance of less than 1 km is predicted. This testifies to the necessity of taking measures to increase safety of launch vehicle flight through observed space debris cluster. In order to increase Ukrainian launch vehicles miss ion safety in the conditions of near space pollution, it is proposed to create the system of pre -flight space analysis, whose tasks are periodic analysis of space situation not less than once in a day, revealing of hazardous approaches, determination of their parameters, and preparation of data to make decision on launch time.

Key words: method of launch time planning, safety of flight through space debris cluster

Bibliography:

1. ESA Operations. For the first time ever, ESA has performed a ‘collision avoidance manoeuvre’ to protect one of its satellites from colliding with a ‘mega constellation’. Electronic resource. – Access mode: https://twitter.com/esaoperations/status/ 1168533241873260544 (Access date 12.09.2019).

2. Klinkrad H. Space Debris – Models and Risk Analysis. Chichester, UK: Praxis Publishing Ltd, 2006. 430 p.

3. Johnson N. L. Orbital Debris: The Growing Threat to Space Operations / Advances in the Astronautical Sciences. 2010. Vol. 137. P. 3-11.

4. Orbital Debris. A Technical Assessment. Washington, D.C.: National Academy Press, 1995. 210 p.

5. Bandyopadhyay P., Sharma R.K., Adimurthy V. Space debris proximity analysis in powered and orbital phases during satellite launch / Advances in Space Research. 2004. Vol. 34. P. 1125-1129. https://doi.org/10.1016/j.asr.2003.10.043

6. Adimurthy V., Ganeshan A. S. Space debris mitigation measures in India / Acta Astronautica. 2005. Vol. 58. P. 168-174. https://doi.org/10.1016/j.actaastro.2005.09.002

7. Schultz E. D., Schultz E. D., Wilde P. D. Mitigation of Collision Hazard for the International Space Station from Globally Launched Objects / 6th IAASS Conference Safety is Not an Option. 21-23 May 2013. Montreal, Canada. Electronic resource. Access mode: https://iaassconference2013.-space-safety.org/ wp-content/uploads/sites/-19/2013/06/ 1420_Shultz.pdf (Access date 12.09.2019).

8. Brevdik G. D., Strub J. E. Determination of acceptable launch windows for satellite collision avoidance / AAS/AIAA Astrodyna-mics Conference. 19-21 August 1991 Pt1. Durango USA. Astrodynamics. P. 345-356.

9. Hejduk M. D., Plakalovic D., New-man L. K., Ollivierre J. C., Hametz M. E., Beaver B. A., Thompson R. C. Trajectory Error and Covariance Realism for Launch Cola Operations / Advances in the Astronautical Sciences. 2013. Vol. 148. P. 2371-2390.

10. Hejduk M. D., Plakalovic D., New-man L. K., Ollivierre J. C., Hametz M. E., Beaver B. A., Thompson R. C. Recommended Risk Assessment Techniques and Thresholds for Launch Cola Operations / Advances in the Astronautical Sciences. 2014. Vol. 150. P. 3061-3080.

11. Handschuh D. A., Wang C., Vidal B. Operational Feedback on Four Years of Collision Risk Avoidance at Launch in Europe / 7th IAASS Conference Space Safety is No Accident, 20-22 October 2014. Fredrichschafen, Germany. P. 355-363. https://doi.org/10.1007/978-3-319-15982-9_42

12. Ihdalov I. М., Kuchma L. D., Poliakov N. V., Sheptun Yu. D. Dinamicheskoe proektirovanie raket. Zadachi dinamiki raket i ikh kosmicheskikh stupenei: mohografiia / pod red. akad. S. N. Koniukhova. Dnepropetrovsk, 2010. 264 s.

13. NIMA TR 8350.2. Department of Defense world geodetic system 1984: Its definition and relationships with local geodetic systems. 3-d ed. National Geospatial-Intelligence Agency, 2000. 174 p.

14. NGA EGM2008 – WGS 84 version. Electronic resource. Access mode to page: http://earth-info.nga.mil/GandG/ wgs84/gravitymod/egm2008/ gm08_wgs84.html. (Access date 12.09.2019).

15. Holubek А. V. Sblizheniie rakety-nositelia s katalogizirovannymi kosmicheskimi ob’ektami v processe vyvedeniia na orbity s nizkim nakloneniem / Izvestiia vysshikh uchebnykh zavadenii. Mashinostroenie. 2018. №2 (695). S. 86-98.

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5.1.2020 Strength and stability of inhomogeneous structures of space technology, consid-ering plasticity and creep

Вацлав Самусевич — Wed, 13 Sep 2023 06:15:53 +0000

5. Strength and stability of inhomogeneous structures of space technology, consid-ering plasticity and creep

Authors:

Hudramovych V. S.², Sirenko V. M.¹, Klimenko D. V.¹, Hart Ye. L.³

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine¹; The Institute of Technical Mechanics, Dnipro, Ukraine²; Oles Honchar Dnipro National University, Dnipro, Ukraine³

Page: Kosm. teh. Raket. vooruž. 2020, (1); 44-56

DOI: https://doi.org/10.33136/stma2020.01.044

Language: Russian

Annotation: The shell structures widely used in space rocket hardware feature, along with decided advantage in the form of optimal combination of mass and strength, inhomogeneities of different nature: structural (different thicknesses, availability of reinforcements, cuts-holes et al.) and technological (presence of defects arising in manufacturing process or during storage, transportation and unforseen thermomechanical effects). The above factors are concentrators of stress and strain state and can lead to early destruction of structural elements. Their different parts are deformed according to their program and are characterized by different levels of stress and strain state. Taking into consideration plasticity and creeping of material, to determine stress and strain state, the approach is effective where the calculation is divided into phases; in each phase the parameters are entered that characterize the deformations of plasticity and creeping: additional loads in the equations of equilibrium or in boundary conditions, additional deformations or variable parameters of elasticity (elasticity modulus and Poisson ratio). Then the schemes of successive approximations are constructed: in each phase, the problem of elasticity theory is solved with entering of the above parameters. The problems of determining the lifetime of space launch vehicles and launching facilities should be noted separately, as it is connected with damages that arise at alternating-sign thermomechanical loads of high intensity. The main approach in lifetime determination is one that is based on the theory of low-cycle and high-cycle fatigue. Plasticity and creeping of material are the fundamental factors in lifetime substantiation. The article deals with various aspects of solving the problem of strength and stability of space rocket objects with consideration for the impact of plasticity and creeping deformations.

Key words: shell structures, stress and strain state, structural and technological inhomogeneity, thermomechanical loads, low-cycle and high-cycle fatigue, lifetime

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20.2.2018 The Use of Special Devices during Launch Pad Development Testing

Вацлав Самусевич — Thu, 07 Sep 2023 12:27:24 +0000

20. The Use of Special Devices during Launch Pad Development Testing

Authors:

Fedenko B. R., Onischenko S. M., Yerofeyev S. I.

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine

Page: Kosm. teh. Raket. vooruž. 2018 (2); 173-177

DOI: https://doi.org/10.33136/stma2018.02.173

Language: Russian

Annotation: One of the tasks of the development tests conducted on a launch pad is verification of its strength properties. The tests are carried out after the launch pad was manufactured and assembled on-site as well as during the whole operating period (if necessary). Load mode was chosen in consideration of cost and possibility of providing the required loading conditions. Two modes of creating the required test load were examined: usage of weights with corresponding mass (load simulators) or special devises (which have smaller mass as compared with load simulators). The descriptions, basic characteristics, advantages and disadvantages of composite and bulk weights and pad loading device are given. This article studies the pad loading device under development. This device enables to conduct static nondestructive tests on the launch pad in order to check its strength after manufacturing and during the whole operating period. The device consists of the load-bearing frame, hydraulic system, locks, control system and measurement system. Advantages of the pad loading device include low materials consumption, low cost in comparison with composite weights (with large load values), provision of the required modes for applying and removing the test load, controlled separate loading of each support of the launch pad, high mobility, short duration of testing, possibility of using launch pads of other rocket complexes with lower or equal test load values for testing. Therefore, the pad loading device enables to achieve the required test load values while having considerably smaller dimensions and mass as compared with composite weights and bigger functional possibilities as compared with bulk weights. Small overall dimensions and operability reduce the number of needed personnel and equipment.

Key words: weight for testing, test load, loading device

Bibliography:

1. ISO 14625:2007. Space systems. Ground support equipment for use at launch, landing or retrieval sites. General requirement. Brought in 01.11.2007. 32 p.

2. Launch Vehicle Mass Dummy: Patent RU2491211 RF: MPK B64G 5/00, B64G 7/00, F42B 15/00 / Dneprotyazhmash. Published 27.08.2013. 12 p.

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4. ISO 16290:2013. Space systems. Definition of the Technology Readiness Levels (TRLs) and their criteria of assessment. Brought in 14.10.2013. 20 p.

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13.1.2018 On Selection of Materials for Creation of Modern LV Thermostating System Mating Hoses

Вацлав Самусевич — Tue, 05 Sep 2023 06:52:56 +0000

13. On Selection of Materials for Creation of Modern LV Thermostating System Mating Hoses

Authors:

Bigun S. O.¹, Yevchik V. S.², Khorolskyi M. S.²

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine¹; State Enterprise DINTEM Ukrainian Research Design-Technological Institute of Elastomer Materials and Products²

Page: Kosm. teh. Raket. vooruž. 2018 (1); 72-84

DOI: https://doi.org/10.33136/stma2018.01.072

Language: Russian

Annotation: A series of materials is proposed for creation of space launch vehicle low-pressure air thermostating systems joints hoses. The topical issues are considered of materials designing with consideration for specific features of the hoses as special industrial rubber articles of launch vehicle launch sites.

Key words:

Bibliography:

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2. Svitlichna R. F., Lotakov V. S., Chumicheva N. P. State and Prospects of Using Rubbers of New Generation in Rubber Industry of Ukraine: Scientific-Technical Report. К., 2001. No. 3. 13 p.

3. Nesterova L. A., Reznichenko S. V., Noskova L. F. et al. Experience of Using BNKS Paraffinate Nitrile Rubber in Formulations of Oil-Resistant Rubbers of Various Purpose at JSC “Uralsky Zavod RTI”. Med. Conf. on rubber: Collection of abstracts. М., 2000. No. 4. 121 p.

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5. Raw Stuff and Materials: Inf. Bull. М., 1999. No. 5. 55 p.

6. Raw Stuff and Materials: Inf. Bull. М., 2001. No. 3. 90 p.

7. Raw Stuff and Materials: Inf. Bull. М., 2001. No. 3. 96 p.

8. Raw Stuff and Materials: Inf. Bull. М., 2000. No. 3. 43 p.

9. Lotakov V. S., Yevchik V. S., Utlenko E. V. et al. Investigation of Operability of Rubbers with Adhesion Additives in Rubber-Metal Valves. Manufacture of Tires, Rubber Products and ATI. М., 1980. No. 4. P. 43-44.

10. Lotakov V. S., Yevchik V. S., Markova L. A. et al. Investigation of Alkali Impact on Adhesive Properties of Ethylene-Propylene Vulcanizing Agents. Caoutchouc and Rubber: Scientific–Technical Report. UNIKTI-DINTEM SE. 1981. No. 6. P. 18-19.

11. Svitlichna R. F., Bogutska E. O., Lotaakov V. S. et al. Technical Carbon of N Series. Prospects of Using in Rubber Mixtures of Caoutchoucs of New Generation: Scientific–Technical Report. К., 2006. No. 3. P. 17-20.

12. Yevchik V. S., Bogutskaya E. A., Khorolsky M. S. Investigations to Select Optimal Options of Replacing Raw Materials and Rubbers with Specifying Guaranteed Service and Storage Life of Rubber Products Being Components of 11K77 Article: Scientific–Technical Report DO-468-2000, UNIKTI-DINTEM SE. 2000. 55 p.

13. Nudelman Z. N., Lavrova L. N. Effective Vulcanization of Fluorine Rubbers. The III Ukr. International Scientific-Technical Conference of Rubber Industry Workers: Collection of abstracts. Dnepropetrovsk, 2000. 43 p.

14. Semyonov G. D., Yevchik V. S., Zaitseva T. P., Lotakov V. S. Prospects of Using New Vulcanizing Systems in Rubber Mixtures Based on Fluoroelastomers: Scientific–Technical Report. К., 2001. No. 3. 18 p.

15. Yevchik V. S., Zaitseva T. P., Khorolsky M. S. Investigations of Physical-Mechanical Characteristics of Rubbers Based on Caoutchoucs of New Generation: Scientific–Technical Report DO-387-89, DF VNIIEMI. Dnepropetrovsk, 2000. 61 p.

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20. Degteva T. G. et al. The Impact of Additives on Thermal Ageing of Rubbers and Model Gaskets Made of SKEP. Caoutchouc and Rubber. М., 1984. No. 8. P. 17-19.

21. Lepetov V. A. Rubber Products. L., 1976. 440 p.

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24. Kornev A. E. et al. Technology of Elastomer Materials. М., 2009. 504 p.

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20.2.2017 Research Support for Development of Launch Vehicle Payload Unit Composite Load-Bearing Compartments

Виктория Лемех — Wed, 09 Aug 2023 12:26:27 +0000

20. Research Support for Development of Launch Vehicle Payload Unit Composite Load-Bearing Compartments

Authors:

Potapov O. М.¹, Kovalenko V. O.¹, Kondrat’ev А. V.², Gaidachuk V. Е.²

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine¹; Kharkiv Aviation Institute, Kharkiv, Ukraine²

Page: Kosm. teh. Raket. vooruž. 2017 (2); 112-120

Language: Russian

Annotation: Some main results of scientific support of development of launch vehicle head module composite loadbearing bays are presented. The methodology is proposed for developing these units. By the example of payload fairing and interstage bay of Cyclone-4 launch vehicle, high efficiency is shown of proposed methodology implementation when selecting their rational design and technological parameters.

Key words:

Bibliography:

1. Degtyarev A. V. Rocket Technology. Problems and Prospects. Selected scientific-technical publications. Dnepropetrovsk, 2014. 420 p.

2. Kovalenko V. A., Kondrat’yev A. V. Use of Polymer Composite Materials in Space Rockets as Reserve of Increasing their Mass and Functional Effectiveness. Aerospace Engineering and Technology. 2011. No. 5 (82). P. 14-20.

3. Kondrat’yev A. V. et al. Analysis of Nomenclature of Type Composite Units of Space Rockets and Structural Schemes Applied for them / A. V. Kondrat’yev, A. G. Dmitrenko, K. D. Stenile, А. А. Tsaritsynsky. Problems of Designing and Manufacturing Flying Vehicle Structures: Collection of scientific works of N. E. Zhukovsky Aerospace University “KhAI”. Issue 3 (79). Kharkiv, 2014. P. 19 – 30.

4. Potapov A. M. et al. Comparison of Payload Fairings of Existing and Prospective Domestic Launch Vehicles and their Foreign Analogs / А. М. Potapov, V. A. Kovalenko, A. V. Kondrat’yev. Aerospace Engineering and Technology. 2015. No. 1(118). P. 35 – 43.

5. Gaidachuk A. V. et al. Methodology of Developing Effective Design and Technological Solutions of Space Rocketry Composite Units: Monography in 2 volumes. Vol. 2. Synthesis of Space Rocketry Composite Units Parameters at Heterogeneous Loading / A. V. Gaidachuk, V. E. Gaidachuk, A. V. Kondrat’yev, V. A. Kovalenko, V. V. Kirichenko, А. M. Potapov / Under the editorship of A. V. Gaidachuk. Kharkiv, 2016. 250 p.

6. Gaidachuk A. V. et al. Methodology of Developing Effective Design and Technological Solutions of Space Rocketry Composite Units: Monography in 2 volumes. Vol. 1. Creation of Space Rocketry Units with Specified Quality of Polymer Composite Materials / A. V. Gaidachuk, V. E. Gaidachuk, A. V. Kondrat’yev, V. A. Kovalenko, V. V. Kirichenko, А. M. Potapov / Under the editorship of A. V. Gaidachuk. Kharkiv, 2016. 263 p.

7. Smerdov A. A. Development of Methods to Design Space Rocketry Composite Materials and Structures: Dissertation of Doctor of Engineering Science: 05.07.02, 05.02.01. М., 2007. 410 p.

8. Slyvyns’kyy V. et al. Basic parameters’ optimization concept for composite nose fairings of launchers / V. Slyvyns’kyy, V. Gajdachuk, V. Kirichenko, A. Kondratiev. 62nd International Astronautical Congress, IAC 2011 (Cape Town, 3-7 October 2011). Red Hook, NY: Curran, 2012. Vol. 9. P. 5701-5710.

9. Gaidachuk V. E. et al. Optimization of Cyclone-4 Launch Vehicle Payload Fairing Design Parameters / V. E. Gaidachuk, V. I. Slivinsky, A. V. Kondrat’yev, A. P. Kushnar’ov, Effectiveness of Honeycomb Structures in Aerospace Products: Proceedings of III International Scientific-Practical Conference (Dnepropetrovsk, 27-29 May 2009). Dnepropetrovsk, 2009. P. 88 – 95.

10. Zinov’yev A. M. et al. Design and Technological Solution and Carrying Capacity of Cyclone-4 Launch Vehicle Interstage Bay Made of Polymer Composite Materials / А. М. Zinov’yev, А. P. Kushnar’ov, A. V. Kondrat’yev, А. М. Potapov, А. P. Kuznetsov, V. A. Kovalenko. Aerospace Engineering and Technology. 2013. No. 3 (100). P. 46-53.

11. Karpov Y. S. Connection of Parts and Units Made of Composite Materials: Monography. Kharkiv, 2006. 359 p.

12. Kondrat’yev A. V. Mass Optimization of Launch Vehicle Payload Fairing Irregular Zones. Problems of Designing and Manufacturing Flying Vehicle Structures: Collection of scientific works of N. E. Zhukovsky Aerospace University “KhAI”. Issue 47 (4). Kharkiv, 2006. P. 126 – 133.

13. Degtyarev A. V. et al. Evaluation of Carrying Capacity of Launch Vehicle Bays Separation System Composite Fitting / A. V. Degtyarev, A. P. Kushnar’ov, V. V. Gavrilko, V. A. Kovalenko, А. V. Kondrat’yev, А. М. Potapov. Space Technology. Missile Armaments: Collection of scientific-technical articles. 2013. Issue 1. P. 18-21.

14. Patent 81537 UA, MPK (2013.01) F42B 15/36 (2006.01) B64D 1/00 Fitting of Rocket’s Three-Layer Shell / О. М. Zinov’yev, О. P. Kuznetsov, V. V. Gavrilko, О. М. Potapov, V. O. Kovalenko et al.; Applicant and patent holder NVF Dniprotechservice, Yuzhnoye SDO. No. u 2012 11210; Claimed 27.09.2012; Published 10.07.13, Bulletin 13. 4 p.

15. Zinov’yev A. M. et al. Manufacturing Technology of Cyclone-4 Launch Vehicle Experimental Large-Sized Interstage Bay Made of Carbon Plastics / А. M. Zinov’yev, А. P. Kushnar’ov, А. V. Kondrat’yev, А. М. Potapov, А. P. Kuznetsov, V. A. Kovalenko. Problems of Designing and Manufacturing Flying Vehicle Structures: Collection of scientific works of N. E. Zhukovsky Aerospace University “KhAI”. Issue 2 (74). Kharkiv, 2013. P. 7 – 17.

16. Zinov’yev A. M. et al. Static Tests of Cyclone-4 Launch Vehicle Experimental Interstage Bay Made of Carbon Plastic / А. М. Zinov’yev, А. P. Kushnar’ov, А. V. Kondrat’yev, А. М. Potapov, А. P. Kuznetsov, V. A. Kovalenko. Aerospace Engineering and Technology. 2013. No. 4(101). P. 28-35.

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