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Heat-resistant MoSi2–NbSi2 and Cr–Ni coatings for rocket engine combustion chambers and respective vacuum-arc deposition technology/ 74th International Astronautical Congress (IAC-23-C2.4.2), Baku, Azerbaijan, 2-6 October 2023.
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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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13. Nadtoka V., Kraiev M., Borisenko A., Bondar D., Gusarova I. Heat-resistant MoSi2–NbSi2 and Cr–Ni coatings for rocket engine combustion chambers and respective vacuum-arc deposition technology/ 74th International Astronautical Congress (IAC-23-C2.4.2), Baku, Azerbaijan, 2-6 October 2023.
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12.1.2024 Hardening of steels modifying their surfaces with ion-plasma nitriding in glow discharge
12.1.2024 Hardening of steels modifying their surfaces with ion-plasma nitriding in glow discharge
12.1.2024 Hardening of steels modifying their surfaces with ion-plasma nitriding in glow discharge

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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
All the models showed convergence with the experimental SRM test data sufficient for engineering estimate (in particular, for maximum chamber pressure and combustion time).
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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

Bibliography:
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  5. 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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  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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]]> 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
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:

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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.
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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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]]> 3.2.2018 Possible Ways of Modernization of VEGA Launch Vehicle AVUM Stage Main Engine Assembly https://journal.yuzhnoye.com/content_2018_2-en/annot_3_2_2018-en/ Thu, 07 Sep 2023 08:42:19 +0000 https://journal.yuzhnoye.com/?page_id=30733
2018 (2); 16-24 DOI: https://doi.org/10.33136/stma2018.02.016 Language: Russian Annotation: The Ukrainian companies Yuzhnoye SDO and SE PA YMZ supply VG143 main engine assembly for Vega LV AVUM upper stage, which is a one-chamber LRE of 250 kg thrust with five ignitions in flight. Key words: main engine assembly , liquid rocket engine , ways of modernization , engine chamber Bibliography: 1. main engine assembly , liquid rocket engine , ways of modernization , engine chamber .
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3. Possible Ways of Modernization of VEGA Launch Vehicle AVUM Stage Main Engine Assembly

Authors:

Prokopchuk O. O., Shulga V. A., Strelchenko E. V., Konokh V. I., Kovalenko A. M., Dibrivny O. V., Lapin O. V., Kukhta A. S.

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine

Page: Kosm. teh. Raket. vooruž. 2018 (2); 16-24

DOI: https://doi.org/10.33136/stma2018.02.016

Language: Russian

Annotation: The Ukrainian companies Yuzhnoye SDO and SE PA YMZ supply VG143 main engine assembly for Vega LV AVUM upper stage, which is a one-chamber LRE of 250 kg thrust with five ignitions in flight. By the present, 11 successful launches of Vega LV have been made. In the process of flight operation, there were no critical comments on engines operation. This LRE has a combination of attractive characteristics, such as high specific pulse, low mass, multiple ignitions in flight, high reliability confirmed by good results of flight test of the prototype engines. The reserve of this engine from the viewpoint of further modernization is far from being exhausted. Enhancing the capabilities of payload injection by launch vehicles into various orbits of artificial Earth satellites is the main task for the developers of ILV as a whole and for the developers of separate assemblies and systems, such as LRE being part of ILV. With consideration for the experience of prototype engines testing, we should note the following ways of main engine assembly modernization: – increasing the specific pulse due to the increase of nozzle expansion ratio; – decreasing the volume of internal manifolds and mass of chamber; – increasing the operation time; – increasing the ignitions number; – increasing the duration of pauses between ignitions and orbital functioning time. Increasing the thrust and specific pulse of Vega LV VG143 main engine assembly and AVUM stage takes place due to the use of pneumopump propellant feeding system instead of standard pressure feeding. Besides, the information is presented on RD859, RD864, RD866 and RD869 prototype engines, the data on their basic characteristics, testing and operation. The below information is of interest to LRE and LV developers.

Key words: main engine assembly, liquid rocket engine, ways of modernization, engine chamber

Bibliography:
1. Shnyakin V., Shul’ga V., Zhivotov A., Dibrivny A. Creating a new generation of space-craft liquid rocket engines basing on pneumopump propellant supply systems. Space Propulsion: International Conference. France, Bordeaux. 2012.
2. Shul’ga V. Development status and improvement methods for upper stage engines of Vega and Cyclone launch vehicles. Space Propulsion; International Conference. Germany, Cologne. 2014.
3. De Rose L., Parmigiani P., Shnyakin V., Shulga V., Pereverzyev V., Caramelli F. Main engine of the Vega fourth stage: characteristics and heritage. 4th International Conference on Launcher Technology “Space Launcher Liquid Propulsion”. Netherlands, Noordwijk. 2018.
4. Kovalenko A. N., Pereverzev V. G., Marchan R. A., Blishun Y. V. Experimental Confirmation of Feasibility of Improving Power-Mass Characteristics of LRE for Vega Launch Vehicle Upper Stage: Paper presentation at the International Scientific-Technical Conference. S. P. Korolev SGAU, 2014.

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3.2.2018 Possible Ways of Modernization of VEGA Launch Vehicle AVUM Stage Main Engine Assembly
3.2.2018 Possible Ways of Modernization of VEGA Launch Vehicle AVUM Stage Main Engine Assembly
3.2.2018 Possible Ways of Modernization of VEGA Launch Vehicle AVUM Stage Main Engine Assembly

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In this case, RD809M engine is RD8 version with tight integration and RD809K engine is its one-chamber version; RD805 engine operating on liquid oxygen + kerosene created on the basis of combustion chamber of RD8 serial control engine of Zenit launch vehicle second stage.; RD835 engine operating on liquid oxygen + kerosene created for the second stages of launch vehicles of Mayak type; the engines and propulsion systems operating on storable propellants, such as RD861K (main engine of Cyclone-4 third stage and Cyclone-4M launch vehicle second stage), DU802 (liquid propulsion system of Krechet autonomous space tug of conversion Dnepr launch vehicle), RD840 (apogee liquid rocket engine of liquid propulsion system of geostationary communication satellite bus), VG143 (main engine assembly of the fourth stage of European Vega launch vehicle), RD864 and RD869 (main engines of Dnepr launch vehicle upper stages).
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2. Yuzhnoye SDO-Developed Upper Stage Liquid Rocket Engines

Authors:

Prokopchuk O. O., Shulga V. A., Strelchenko E. V., Dibrivny O. V., Kukhta A. S.

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine

Page: Kosm. teh. Raket. vooruž. 2018 (2); 8-15

DOI: https://doi.org/10.33136/stma2018.02.008

Language: Russian

Annotation: One of the important directions in the development of Yuzhnoye SDO liquid rocket engines is creation of the engines for launch vehicle upper stages, boosters, space tugs and takeoff-and-landing vehicles. The article presents an overview of Yuzhnoye SDO – developed liquid rocket engines, their basic characteristics, distinctive features and the current status of development and operation. The article presents the information on the following engines: RD858 and RD859 operating on storable propellants, for lunar takeoff-and-landing Block E module; RD809M and RD809K operating on liquid oxygen + kerosene created on the basis of RD8 serial control engine of Zenit launch vehicle second stage. In this case, RD809M engine is RD8 version with tight integration and RD809K engine is its one-chamber version; RD805 engine operating on liquid oxygen + kerosene created on the basis of combustion chamber of RD8 serial control engine of Zenit launch vehicle second stage.; RD835 engine operating on liquid oxygen + kerosene created for the second stages of launch vehicles of Mayak type; the engines and propulsion systems operating on storable propellants, such as RD861K (main engine of Cyclone-4 third stage and Cyclone-4M launch vehicle second stage), DU802 (liquid propulsion system of Krechet autonomous space tug of conversion Dnepr launch vehicle), RD840 (apogee liquid rocket engine of liquid propulsion system of geostationary communication satellite bus), VG143 (main engine assembly of the fourth stage of European Vega launch vehicle), RD864 and RD869 (main engines of Dnepr launch vehicle upper stages). The information presented in the article is of interest to liquid rocket engines and launch vehicles developers.

Key words: main engine, engine development test, takeoff-and-landing module, pneumatic pump unit

Bibliography:
1. Liquid Rocket Engines, Propulsion Systems, Onboard Power Sources Developed by Propulsion Systems Design Office of Yuzhnoye SDO / Under scientific editorship of S. N. Konyukhov, Academician of NAS of Ukraine, V. N. Shnyakin, Candidate of Engineering Science. Dnepropetrovsk, 2008. 466 p.
2. Liquid Rocket Engines. Description and Basic Technical Data / Under scientific editorship of S. N. Konyukhov, Academician of NAS of Ukraine, V. N. Shnyakin, Candidate of Engineering Science. Dnepropetrovsk, 1996. 84 p.
3. Prokopchuk A. A. et al. New Possibilities for Creation of Apogee Propulsion Systems with Pneumopump Propellant Supply System. Paper presentation at Conference “Space Propulsion”, 2018, Spain.
4. Shnyakin V. N., Shulga V. A., Dibrivny A. V. Possibilities of Creating New LRE Based on Mature Technologies. Space Technology. Missile Armaments: Collection of scientific-technical articles. 2011. Issue 2. P. 61-71.

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2.2.2018 Yuzhnoye SDO-Developed Upper Stage Liquid Rocket Engines
2.2.2018 Yuzhnoye SDO-Developed Upper Stage Liquid Rocket Engines
2.2.2018 Yuzhnoye SDO-Developed Upper Stage Liquid Rocket Engines

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]]> 17.1.2019 Development of Prospective Small-Size Auxiliary SMR of New Type https://journal.yuzhnoye.com/content_2019_1-en/annot_17_1_2019-en/ Wed, 24 May 2023 16:00:35 +0000 https://journal.yuzhnoye.com/?page_id=27722
Testing SRE with pyroxiline powder grain showed that the optimum design of the engine can be developed only with the application of the specially developed design procedure of the gas-dynamic flow pattern of powder gases in the engine chamber with definition of field of pressure and velocity.
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17. Development of Prospective Small-Size Auxiliary SMR of New Type

Authors:

Toloch’yants G. E., Mikhailov M. S., Magdin E. K., Oglikh V. V.

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine

Page: Kosm. teh. Raket. vooruž. 2019, (1); 114-121

DOI: https://doi.org/10.33136/stma2019.01.114

Language: Russian

Annotation: This article considers essentially new versions of small-sized solid propellant rocket engines (SRE), designed for rocket and spacecraft flight control with serial artillery pyroxiline powder taken as grain and solidpropellant gas generators discretely operating into the receiver. Preliminary results of design and experimental activities, performed in Yuzhnoye SDO, showed the possibility in principle and practicability to develop two new types of advanced small-sized SRE. Testing SRE with pyroxiline powder grain showed that the optimum design of the engine can be developed only with the application of the specially developed design procedure of the gas-dynamic flow pattern of powder gases in the engine chamber with definition of field of pressure and velocity. Such procedure has been developed based on Ansys software package. The article describes areas of further design and experimental activities, fulfilment of which will provide development of production models of the described engines. Intraballistic characteristics design procedure, mentioned in the article, can be used to design new type of micropulse SRE with less than 0.1 s burn time. This article will also facilitate definition of the application area for discrete solid-propellant propulsion systems, where they get the edge over the cold gas gas-jet systems.

Key words: procedure, microSRE, gas-jet system, heat-transfer factor

Bibliography:

1. Kovalenko N. D., Kukushkin V. I. Triumph I tragediya systemy upravleniya vektorom tyagi dvigatelya ZD65 vduvom kamernogo gaza v soplo// Kosmicheskaya technika. Raketnoe vooruzhenie: Sb. nauch.-techn. st. 2014. Vyp. 1. Dnepropetrovsk: GP KB «Yuzhnoye». P. 97-106.
2. Oglykh V. V., Vakhromov V. A., Kirichenko A. S., Kosenko M. G. Razrabotka porokhovykh accumulyatorov davlenia dlya minometnogo starta raket – vazhneishee uslovie ego uspeshnoy realizatsii / Kosmicheskaya technika. Raketnoe vooruzhenie: Sb. nauch.-techn. st. 2016. Vyp. 1. Dnepropetrovsk: GP KB «Yuzhnoye». P. 88-92.
3. Golubev K. S., Svetlov V. G. Proektirovanie zenitnykh upravlyaemykh raket. M.: Izd-vo MAH, 2001. 730 p.
4. Oglykh V. V., Tolochyants G. E., Mikhailov N. S., PopkovV. N. Eksperimentalnye issledovania vozmozhnosti sozdania impulsnogo RDTT s malym vremenem raboty/ Kosmicheskaya technika. Raketnoe vooruzhenie: Sb. nauchn.-techn. st. 2016. Vyp. 2. Dnepr: GP KB «Yuzhnoye». P. 30-34.
5. Belyaev N. M., Belik N. P., Uvarov Ye. I. Reaktyvnye systemy upravleniya kosmicheskykh letatelnykh apparatov. M.: Mashinostroenie, 1979. 232 p.
6. Gubertov A. M., Mironov V. V., Borisov D. M. Gazodynamicheskie i teplophysicheskie process v raketnykh dvigatelyakh na tverdom toplive. M.: Mashinostroenie, 2004.
7. Kutateladze S. S. Teploperedacha i hydrodynamicheskoe soprotivlenie. Energoatomizdat, 1990. 368 p.
8. Scherbakov M. A. Opredelenie coeffitsientov teplootdachi pri modelirovanii zadach v Ansys CFX // Dvigateli i energoustanovki aerokosmicheskykh letatelnykh apparatov: Sb. nauch. statey. M.: Nauch.- techn. Centr im. A. Lyulki, 2014.
9. Moskvichev A. V. Primenimost’ modeley turbulentnosti, realizovannykh v Ansys CFX dlya issledovaniya gasodynamiki v schelevom kanale TNA ZhRD. Voronezhskiy gosudarstvenniy technicheskiy universitet, 2015.
10. Magdin E. K., Oglykh V. V., Rozlivan A. B. Tverdotoplivnaya dvigatelnaya ustanovka orientatsii I stabilizatsii descretnogo deistviya dlya upravleniya kosmicheskimi obiektami / Vestn. dvigatelestroiteley. 2017. Vyp. 2. P. 108-111.

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17.1.2019 Development of Prospective Small-Size Auxiliary SMR of New Type
17.1.2019 Development of Prospective Small-Size Auxiliary SMR of New Type
17.1.2019 Development of Prospective Small-Size Auxiliary SMR of New Type

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]]> 12.2.2019 Procedure of acoustic loads measuring during the ILV launch https://journal.yuzhnoye.com/content_2019_2-en/annot_12_2_2019-en/ Mon, 15 May 2023 15:46:04 +0000 https://journal.yuzhnoye.com/?page_id=27214
2019, (2); 92-95 DOI: https://doi.org/10.33136/stma2019.02.092 Language: Russian Annotation: The paper considers the main aspects of acoustic loads measuring during the rocket launch as well as problem of staff, equipment and environment protection from destructive effect of noise, generated by rocket’s engine. As an example of use of this procedure, results of calculation of noise levels, provided by the sound attenuating chamber, and measurement data during the pulsed wind tunnel tests have been presented. As an outcome of calculations and measurements, done by the proposed procedure, outlines of the noise-safe zone were successfully defined and number of modifications suggested for the sound attenuating chamber to reduce the acoustic loads it generates.
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12. Procedure of acoustic loads measuring during the ILV launch

Authors:

Batutina T. Y., Bondar’ D. S.

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine

Page: Kosm. teh. Raket. vooruž. 2019, (2); 92-95

DOI: https://doi.org/10.33136/stma2019.02.092

Language: Russian

Annotation: The paper considers the main aspects of acoustic loads measuring during the rocket launch as well as problem of staff, equipment and environment protection from destructive effect of noise, generated by rocket’s engine. Proposed is the procedure of acoustic loads measurement and calculation of the minimal acoustic noise-safe launching rocket distance, at which it is safe to put equipment and operating personnel. This procedure is based on sharing of the numerical simulation of the exhaust jet in the ANSYS software system, engineering approaches in calculation of propagation and extinction of acoustic waves and measurement of the actual values of acoustic loads with the use of several noise meters. Effective sanitary standards of noise safety were taken to define the duration and power of acoustic loads safe for the personnel. As an example of use of this procedure, results of calculation of noise levels, provided by the sound attenuating chamber, and measurement data during the pulsed wind tunnel tests have been presented. As an outcome of calculations and measurements, done by the proposed procedure, outlines of the noise-safe zone were successfully defined and number of modifications suggested for the sound attenuating chamber to reduce the acoustic loads it generates.

Key words: Supersonic jet, calculation of acoustic loads, acoustic measurements, acoustic protection

Bibliography:
1. Opredelenie summarnogo urovnya shuma neskolkykh istochnikov. URL: http//studbooks.net/39077/bzhd/ opredelit_summarnyy_uroven_neskolkih_ istochnikov_shuma (data obrascheniya: 06.08.2017).
2. DSN 3.3.6.037-99. Sanitarni normy vyrobnychogo shumu, ultrazvuku ta infrazvuku.
3. Khekl M., Muller Kh. A. Spravochnik po tekhnicheskoy akustike. 1980. 438 s.

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12.2.2019 Procedure of acoustic loads measuring during the ILV launch
12.2.2019 Procedure of acoustic loads measuring during the ILV launch
12.2.2019 Procedure of acoustic loads measuring during the ILV launch

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