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Aerospace
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Combustion of Solid Propellants
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Combustion of Solid Propellants

A special issue of Aerospace (ISSN 2226-4310). This special issue belongs to the section "Aeronautics".

Deadline for manuscript submissions: 30 September 2025 | Viewed by 5748

Special Issue Editor

Dr. Wen Ao

E-Mail Website
Guest Editor
National Key Laboratory of Solid Rocket Propulsion, Northwestern Polytechnical University, Xi’an 710072, China
Interests: solid propellant combustion; advanced energy management technologies; metal combustion
Special Issues, Collections and Topics in MDPI journals

Special Issue Information

Dear Colleagues,

Solid rocket motors (SRMs), due to their various advantages, including their inherent simplicity, high reliability, and quick response, play an important role in space launch vehicle design and production. Surveying the demands surrounding the continuous performance improvements and technological development of SRMs, high energy, high pressure, and extreme dimensions are the most important development trends. As combustion is the most critical process for energy release from solid propellants, there has been a growing interest in the fundamental issues inherent to it; indeed, combustion modeling, combustion diagnostics, and two-phase flow simulation are the typical technical challenges faced in efforts to increase the reliability of SRMs. Improving combustion efficiency and burning rate adjustment are also an urgent priority. Building on these needs, this Special Issue aims to provide an overview of the most recent advances in the field of SRM combustion. Potential topics include, but are not limited to, combustion modeling, combustion adjustment, high-pressure combustion, combustion instability assessment, metal behavior and two-phase flow simulation.

Dr. Wen Ao
Guest Editor

Manuscript Submission Information

Manuscripts should be submitted online at www.mdpi.com by registering and logging in to this website. Once you are registered, click here to go to the submission form. Manuscripts can be submitted until the deadline. All submissions that pass pre-check are peer-reviewed. Accepted papers will be published continuously in the journal (as soon as accepted) and will be listed together on the special issue website. Research articles, review articles as well as short communications are invited. For planned papers, a title and short abstract (about 100 words) can be sent to the Editorial Office for announcement on this website.

Submitted manuscripts should not have been published previously, nor be under consideration for publication elsewhere (except conference proceedings papers). All manuscripts are thoroughly refereed through a single-blind peer-review process. A guide for authors and other relevant information for submission of manuscripts is available on the Instructions for Authors page. Aerospace is an international peer-reviewed open access monthly journal published by MDPI.

Please visit the Instructions for Authors page before submitting a manuscript. The Article Processing Charge (APC) for publication in this open access journal is 2400 CHF (Swiss Francs). Submitted papers should be well formatted and use good English. Authors may use MDPI's English editing service prior to publication or during author revisions.

Keywords

  • propellant combustion
  • combustion adjustment
  • combustion modeling
  • high-pressure combustion

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Related Special Issue

Published Papers (4 papers)

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Research

15 pages, 7344 KiB  
Article
The Effect of 0–8 MPa Environmental Pressure on the Ignition and Combustion Process of CL20/NEPE Solid Propellant
by Wenxiang Cai, Wei Li and Zhixiang Wang
Aerospace 2024, 11(8), 672; https://doi.org/10.3390/aerospace11080672 - 15 Aug 2024
Viewed by 1226
Abstract
In order to study the effect of pressure on the ignition and combustion process of CL-20/NEPE solid propellant, the ignition delay, burning rate, and maximum combustion temperature of different solid propellant formulations with an ambient pressure of 0.1~8.0 MPa were measured experimentally by [...] Read more.
In order to study the effect of pressure on the ignition and combustion process of CL-20/NEPE solid propellant, the ignition delay, burning rate, and maximum combustion temperature of different solid propellant formulations with an ambient pressure of 0.1~8.0 MPa were measured experimentally by a solid propellant laser ignition experiment system, and the agglomeration process and the characteristics of condensed phase combustion products were analyzed. The experimental results show that, with the increase of pressure, the ignition-delay time decreases, and the burning rate and the maximum combustion temperature increase. With the increase of pressure, the influence on propellant ignition and combustion characteristics becomes smaller. In the experiment, the dynamic agglomeration phenomenon of aluminum particles in the propellant was recorded by a high-speed camera combined with a microscopic camera lens, and the dynamic agglomeration phenomenon of the combustion surface of the propellant and the dynamic agglomeration phenomenon, after the initial agglomeration was separated from the surface, were analyzed and expounded. Based on the experiment and combined with the agglomeration phenomenon, a mathematical model capable of predicting the particle size of aluminum aggregates was proposed. Full article
(This article belongs to the Special Issue Combustion of Solid Propellants)
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17 pages, 6707 KiB  
Article
Effects of Wire-Wrapping Patterns and Low Temperature on Combustion of Propellant Embedded with Metal Wire
by Qiu Wu, Jiangong Zhao and Quanbin Ren
Aerospace 2024, 11(8), 639; https://doi.org/10.3390/aerospace11080639 - 6 Aug 2024
Viewed by 1443
Abstract
Incorporating silver wires into propellant has emerged as a highly effective strategy for enhancing propellant burning rates, a technique extensively deployed in the construction of numerous fielded sounding rockets and tactical missiles. Our research, employing a multi-faceted approach encompassing thermogravimetric-differential scanning calorimetry measurements [...] Read more.
Incorporating silver wires into propellant has emerged as a highly effective strategy for enhancing propellant burning rates, a technique extensively deployed in the construction of numerous fielded sounding rockets and tactical missiles. Our research, employing a multi-faceted approach encompassing thermogravimetric-differential scanning calorimetry measurements (TG-DSC), combustion diagnoses, burning rate tests, and meticulous collection of condensed combustion products, sought to elucidate how variations in silver wire quantity and winding configuration impact the combustion properties of propellants. Our findings underscore the remarkable efficacy of double tightly twisted silver wire in significantly boosting propellant burning rates under ambient conditions. Moreover, at lower temperatures, the reduced gap between the propellant and silver wire further magnifies the influence of silver wire on burning rates. However, it is noteworthy that the relationship between burning speed and combustion efficiency is not deterministic. While a smaller cone angle of the burning surface contributes to heightened burning rates, it concurrently exacerbates the polymerization effect of vapor phase aluminum particles, consequently diminishing propellant combustion efficiency. Conversely, propellants configured with sparsely twinned silver wires exhibit notable enhancements in combustion efficiency, despite a less pronounced impact on the burning rate attributed to the larger cone angle of the burning surface. Remarkably, these trends persist at lower temperatures. Based on the principle of heat transfer balance, a theoretical model for the combustion of propellants with wire inserts is developed. The reliability of this theoretical model is validated through a comparison of calculated values with experimental data. Our research outcomes carry significant implications for guiding the application and advancement of the silver wire method in solid propellants for solid rocket motors, offering valuable insights to inform future research and development endeavors in this domain. Full article
(This article belongs to the Special Issue Combustion of Solid Propellants)
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20 pages, 11814 KiB  
Article
Design of Reverse Bleed Slot for Curved Axisymmetric Inlet Based on Kantrowitz Criterion and Flow Field Characteristics
by Yongzhou Li, Di Sun, Xinhui Tian, Yiqi Yuan, Xisheng Luo and Kunyuan Zhang
Aerospace 2024, 11(7), 553; https://doi.org/10.3390/aerospace11070553 - 4 Jul 2024
Viewed by 995
Abstract
Conventional forward bleed slots reduce the hypersonic inlet starting Mach number but suffer from excessive flow leakage after restart. This paper proposes a novel reverse bleed slot design method for curved axisymmetric inlets of a solid-fuel scramjet. Leveraging the Kantrowitz criterion and detailed [...] Read more.
Conventional forward bleed slots reduce the hypersonic inlet starting Mach number but suffer from excessive flow leakage after restart. This paper proposes a novel reverse bleed slot design method for curved axisymmetric inlets of a solid-fuel scramjet. Leveraging the Kantrowitz criterion and detailed flow analysis, the method optimizes bleed slot placement, number, area, and angle. Results show superior aerodynamic performance by placing slots in the non-starting region of the internal compression section, considering both unstarted flow and separation bubble dynamics during restart. Each bleed slot area is calculated successively down-stream based on the Kantrowitz criterion. Finally, the effects of bleed slot angle have been extensively studied. The key inlet performance reaches its optimum at a slot angle of approximately 130°, achieving a significant reduction in the starting Mach number (from 4.80 to 3.65) and a 50% decrease in bleed flow rate compared to the forward slot design. This method demonstrates its feasibility and effectiveness, enabling substantial improvement in inlet starting performance with minimal flow loss. Full article
(This article belongs to the Special Issue Combustion of Solid Propellants)
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17 pages, 3699 KiB  
Article
Numerical and Experimental Analyses of the Effect of Water Injection on Combustion of Mg-Based Hydroreactive Fuels
by Shiyao Shao, Songchen Yue, Hong Qiao, Peijin Liu and Wen Ao
Aerospace 2024, 11(7), 542; https://doi.org/10.3390/aerospace11070542 - 1 Jul 2024
Cited by 1 | Viewed by 1335
Abstract
The energy release process of the Mg-based hydroreactive fuels directly affects the performance of water ramjet engines, and the burning rate is one of the key parameters of the Mg-based hydroreactive fuels. However, there is not enough in-depth understanding of the combustion process [...] Read more.
The energy release process of the Mg-based hydroreactive fuels directly affects the performance of water ramjet engines, and the burning rate is one of the key parameters of the Mg-based hydroreactive fuels. However, there is not enough in-depth understanding of the combustion process of Mg-based hydroreactive fuels within the chamber of water ramjet engines, and there is a lack of effective means of prediction of the burning rate. Therefore, this paper aims to examine the flame structure of Mg-based hydroreactive fuels with a high metal content and analyze the impact of the water injection velocity and droplet diameter on the combustion property. A combustion experiment system was designed to replicate the combustion of Mg-based hydroreactive fuels within water ramjet engines, and the average linear burning rate was calculated through the target line method. On the basis of the experiment, a combustion–flow coupling solution model of Mg-based hydroreactive fuels was formulated, including the reaction mechanism between Mg/H2O and the decomposition products from an oxidizer and binder. The model was validated through experimental results with Mg-based hydroreactive fuels at various pressures and water injection velocities. The mean absolute percentage error (MAPE) in the experimental results was less than 5%, proving the accuracy and validity of the model. The resulting model was employed for simulating the combustion of Mg-based hydroreactive fuels under different water injection parameters. The addition of water injection resulted in the creation of a new high-temperature region, namely the Mg/H2O non-premixed combustion region in addition to improving the radial diffusion of the flame. With the increasing water injection velocity, the characteristic distance of Mg/H2O non-premixed combustion region is decreased, which enhances the heat transfer to burning surface and accelerates the fuel combustion. The impact of droplet parameters was investigated, revealing that larger droplets enhance the penetration of the fuel-rich gas, which is similar to the effect of injection velocity. However, when the droplet size becomes too large, the aqueous droplets do not fully evaporate, resulting in a slight decrease in the burning rate. These findings enhance the understanding of the mechanisms behind the burning rate variation in Mg-based hydroreactive fuels and offer theoretical guidance for the optimal selection of the engine operating parameters. Full article
(This article belongs to the Special Issue Combustion of Solid Propellants)
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