Search Results for “combustion chamber pressure” – Collected book of scientific-technical articles https://journal.yuzhnoye.com Space technology. Missile armaments Sun, 15 Mar 2026 17:40:21 +0000 en-GB hourly 1 https://journal.yuzhnoye.com/wp-content/uploads/2020/11/logo_1.svg Search Results for “combustion chamber pressure” – Collected book of scientific-technical articles https://journal.yuzhnoye.com 32 32 2.2.2025 Performance analysis and validation of a monopropellant air-detonation ramjet engine https://journal.yuzhnoye.com/content_2025_2-en/annot_2_2_2025-en/ Tue, 27 Jan 2026 08:06:29 +0000 https://journal.yuzhnoye.com/?page_id=35818
The infl uence of diff erent combustion chamber lengths and injector confi gurations was also assessed, revealing that geometric optimization plays a crucial role in maintaining detonation stability across diff erent temperature regimes. The fi ndings contribute to optimizing heat release dynamics and pressure gain within the detonation chamber, off ering valuable insights into designing lightweight, energy-effi cient engines for future aerospace systems.
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2. Performance analysis and validation of a monopropellant air-detonation ramjet engine

Date of receipt of the article for publication: 15.10.2025

Date of acceptance of the article for publication after review: 29.10.2025

Date of publication: 27.01.2026

e-ISSN: 2617-5533

Authors:

Stoliarchuk V. V., Tertyshnyk S. V.

ORCID authors:

Stoliarchuk V. V. ORCID, Tertyshnyk S. V. ORCID

Organization:

Page: Kosm. teh. Raket. vooruž. 2025 (2); 12-23

DOI: https://doi.org/10.33136/stma2025.02.012

Language: English

Annotation: The increasing relevance of alternative propulsion systems necessitates an exploration of the potential of monopropellant detonation engines for compact and effi cient aerospace applications. This study aimed to investigate the operating parameters and performance characteristics of a direct-fl ow air-detonation propulsion system operating on environmentally friendly monopropellants. The research was based on a combination of experimental methods and numerical simulation using validated thermochemical models. It presents the results of a series of tests conducted with modifi ed engine geometries under varying inlet temperature and pressure conditions, focusing on achieving a stable detonation wave and analysing its propagation features. A detailed comparison between experimental pressure data and numerical predictions showed a deviation of less than 6.5 %, validating the reliability of the simulation model for practical applications. The infl uence of diff erent combustion chamber lengths and injector confi gurations was also assessed, revealing that geometric optimization plays a crucial role in maintaining detonation stability across diff erent temperature regimes. The study identifi ed critical fl ow parameters for successful ignition and detonation maintenance without external oxidizers, and highlighted the performance of two promising monopropellant compositions, including a modifi ed pronit-based propellant. The fi ndings contribute to optimizing heat release dynamics and pressure gain within the detonation chamber, off ering valuable insights into designing lightweight, energy-effi cient engines for future aerospace systems. The practical value of this research lies in the potential of applying its results in the design of advanced aerospace propulsion systems that feature compact size and environmental friendliness.

Key words: detonation combustion, wave stability, experimental simulation, thermal dynamics, geometric optimisation

Bibliography:

1. Zhang H., Jiang L., Liu W. D. & Liu S. J. Characteristic of rotating detonation wave in an H2/Air hollow chamber with Laval nozzle. International Journal of Hydrogen Energy. 2021. 46 (24). 13389–13401. https://doi.org/10.1016/j.ijhydene.2021.01.143
2. Xue S., Ying Z., Hu M., & Zhou C. Experimental study on the rotating detonation engine based on a gas mixture. Frontiers in Energy Research. 2023. 11. 1136156. https://doi.org/10.3389/fenrg.2023.1136156
3. Xue S., Ying Z., Ma H., & Zhou C. Experimental investigation on two-phase rotating detonation fueled by kerosene in a hollow directed combustor. Frontiers in Energy Research. 2022. 10, 951177. https://doi.org/10.3389/fenrg.2022.951177
4. Kawalec M., Wolanski P., Perkowski W., & Bilar A. Development of a liquid-propellant rocket powered by a rotating detonation engine. Journal of Propulsion and Power. 2023. 39(4). 554–561. https://doi.org/10.2514/1.B38771
5. Zolotko O. Y., Zolotko O. V., Aksyonov O. S., Stoliarchuk V. V., & Cherniavskyi O. S. Analysis of the characteristics of the ejector regime of the impulse-detonation engine of the combined cycle of acceleration. Aerospace technic and technology. 2024. 6 (200). 52–59. https://doi.org/10.32620/aktt.2024.6.05
6. Camacho, R. G., & Huang, C. Component-based reduced order modelling of two-dimensional rotating detonation engine with non-uniform injection. AIAA SCITECH 2025 Forum.
https://doi.org/10.2514/6.2025-1397
7. Feng W., Zhang Q., Xiao Q., Meng H., Han X., Cao Q., Huang H., Wu B., Xu H., & Weng C. Effects of cavity length on operating characteristics of a ramjet rotating detonation enjine fueled by liquid kerosene. Fuel. 2023. 332. 126129. https://doi.org/10.1016/j.fuel.2022.126129
8. Bennewitz J. W., Bigler B. R., Ross M. C., Danczyk S. A., Hargus W. A. Jr. & Smith R. D. Performance of a rotating detonation rocket engine with various convergent nozzles and chamber lengths. Energies. 2021. 14(8). 2037. https://doi.org/10.3390/en14082037
9. Curran D., Wheatley V. & Smart M. High Mach number operation of accelerator scramjet engine. Journal of Spacecraft and Rockets. 2023. 60(3). https://doi.org/10.2514/1.A35511
10. Sun D., Dai Q., Chai W. S., Fang W. & Meng H. Experimental studies on parametric effects and reaction mechanisms in electrolytic decomposition and ignition of HAN solutions. ACS Omega. 2022. 7(22). 18521–18530. https://doi.org/10.1021/acsomega.2c01183
11. Stoliarchuk V. V. Validation of efficiency enhancement methods for detonation jet engines. Aerospace technic and technology. 2024. 4(1). 82–88. https://doi.org/10.32620/aktt.2024.4sup1.12
12. Wang J., Liu Y., Huang W., Zhang Y. & Qiu H. Direct numerical simulation of inflow boundary-layer turbulence effects on cavity flame stabilisation in a model scramjet combustor. Aerospace Science and Technology. 2025. 165. 110463. https://doi.org/10.1016/j.ast.2025.110463
13. Li W., Oh H. & Ladeinde F. Comparison of flamelet and transported species-based modeling of rotating detonation engines. AIAA SCITECH 2024 Forum. https://doi.org/10.2514/6.2024-2599.
14. Chen Y., Liu S., Peng H., Zhong S., Zhang H., Yuan X., Fan W. & Liu W. Propagation and heat release characteristics of rotating detonation in a ramjet engine with a divergent combustor. Physics of Fluids, 2025 37(2), 026132. https://doi.org/10.1063/5.0254419
15. Kailasanath K. Review of propulsion applications of detonation waves. AIAA Journal. 2000. 38(9). 1698–1708. https://doi.org/ 10.2514/2.1156
16. Heiser W. H., & Pratt D. T. Thermodynamic cycle analysis of pulse detonation engines. Journal of Propulsion and Power. 2002. 18(1), 68–76. https://doi.org/10.2514/2.5899
17. Munipalli R., Shankar V., Wilson D. R., Kim H., Lu F. K. & Liston G. Performance assessment of ejector-augmented pulsed detonation rockets. In 39th Aerospace Sciences Meeting and Exhibit (Paper 2001-0830). Reno: AIAA. https://doi.org/10.2514/6.2001-830
18. Lu F. K. & Braun E. M. Rotating detonation wave propulsion: Experimental challenges, modeling, and engine concepts. Journal of Propulsion and Power. 2014. 30(5). 1125–1142. https://doi.org/10.2514/1.B34802
19. Armbruster W. et al. Design and testing of a hydrogen–oxygen pre-detonator for RDEs. CEAS Space Journal. 2025. 17. 969-979.
https://doi.org/10.1007/s12567-025-00605-y

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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
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).
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8. Theoretic-experimental evaluation of the solid-propellant grain erosive burning

e-ISSN: 2617-5533

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:
  1. Arkhipov V. Erosionnoe gorenie condensirovannykh system. Sb. tr. ІХ Vserossiyskoy nauch. conf. 2016 g. (FPPSM-2016). Tomsk, 2016.
  2. 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
  3. Sabdenov K. , Erzda M., Zarko V. Ye. Priroda i raschet skorosti erozionnogo goreniya tverdogo raketnogo topliva. Inzhenerniy journal: nauka i innovatsii, 2013. Vyp. 4.
  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.
  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
  6. Williams, Forman A., Combustion Theory. The Benjamin/Cummings Publishing , Menlo Park, 1985.
  7. Irov Yu. D., Keil E. V., Maslov B.N., Pavlukhin Yu. A., Porodenko V. V.,
    Stepanov Ye. A. Gasodynamicheskie funktsii. Mashinostroenie, Moskva, 1965.
  8. William Orvis. EXCEL dlya uchenykh, inzhenerov i studentov. Kiev: «Junior», 1999.

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]]> 10.1.2020 Calculation and selection of parameters for a propellant consumption diagram of dual-thrust main SRM https://journal.yuzhnoye.com/content_2020_1-en/annot_10_1_2020-en/ Wed, 12 Jun 2024 14:58:34 +0000 https://journal.yuzhnoye.com/?page_id=31037
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. solid propellant charge mass , propellant combustion rate , combustion chamber pressure , operation time in starting and cruise modes , combustion chamber pressure difference .
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10. Calculation and selection of parameters for a propellant consumption diagram of dual-thrust main SRM

e-ISSN: 2617-5533

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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]]> 9.1.2016 Evaluation of SRM Intraballistic Parameters Scatter by Monte-Carlo Method https://journal.yuzhnoye.com/content_2016_1/annot_9_1_2016-en/ Tue, 23 May 2023 13:04:25 +0000 https://journal.yuzhnoye.com/?page_id=27617
2016 (1); 59-62 Language: Russian Annotation: The purpose of work is development of simple and accurate procedure of calculation of scatter of SRM intraballistic parameters – combustion chamber pressure and combustion products flow rate.
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9. Evaluation of SRM Intraballistic Parameters Scatter by Monte-Carlo Method

e-ISSN: 2617-5533

Authors:

Fomenko V. S., Kalnysh R. V., Popov V. A.

Organization:

Yangel Yuzhnoye State Design Office, Dnipro, Ukraine

Page: Kosm. teh. Raket. vooruž. 2016 (1); 59-62

Language: Russian

Annotation: The purpose of work is development of simple and accurate procedure of calculation of scatter of SRM intraballistic parameters – combustion chamber pressure and combustion products flow rate. The problem is solved by Monte Carlo method (statistic tests) which allows constructing the function of distribution of composition (sum, product) of random values of specified distribution laws. The accuracy of problem solution is determined by quality of pseudo random values generation.

Key words:

Bibliography:

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9.1.2016 Evaluation of SRM Intraballistic Parameters Scatter by Monte-Carlo Method
9.1.2016 Evaluation of SRM Intraballistic Parameters Scatter by Monte-Carlo Method
9.1.2016 Evaluation of SRM Intraballistic Parameters Scatter by Monte-Carlo Method
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