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Combustion Evaluation and Control of Solid Rocket Motors

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

Deadline for manuscript submissions: closed (31 December 2023) | Viewed by 26747

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Dr. Wen Ao

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National Key Laboratory of Solid Rocket Propulsion, Northwestern Polytechnical University, Xi’an 710072, China
Interests: solid propellant combustion; advanced energy management technologies; metal combustion
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Dear Colleagues,

Due to their various advantages, including their inherent simplicity, high reliability, and quick response, solid rocket motors (SRM) play an important role in space launch vehicles. Taking account of demands around the continuous performance improvement and technological development of SRM, high energy, high pressure, high overload, wide adaptation, and extreme dimension are the most important development trends. As combustion is the most critical process for the energy release of solid propellants, there is growing interest in the fundamental issues around it in SRM. Indeed, combustion diagnostics on a mesoscopic scale, condensed phase evolution, and nonlinear combustion instability mechanism are the typical technical challenges to be faced in order to increase the reliability of SRM. Improving combustion efficiency via propellant modification is also an urgent need. Building on this vision, this Special Issue aims to provide an overview of the most recent advances in the field of SRM combustion and control. Potential topics include but are not limited to thermal decomposition characteristics of propellants, aluminum particle behavior, advanced combustion diagnostics methods, chemical analysis of condensed combustion products, combustion instability assessment, and optimization of propellant formulation.

Dr. Wen Ao
Guest Editor

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Keywords

solid propellant combustion
mesoscopic measurement
condensed combustion products
high combustion efficiency

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Related Special Issue

Combustion of Solid Propellants in Aerospace (4 articles)

Published Papers (11 papers)

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Research

15 pages, 3645 KiB

Open AccessArticle

Tuning the Ballistic Performance of a Single-Burning-Rate Grain Solid Rocket Motor via New Discontinuous Embedded Metal Wires

by Qiu Wu and Quanbin Ren

Aerospace 2024, 11(4), 308; https://doi.org/10.3390/aerospace11040308 - 15 Apr 2024

Cited by 1 | Viewed by 1544

Abstract

This work proposes a new effective method to realize variable thrust through discontinuous embedded metal wires in the solid rocket motor (SRM). We aimed to study the influence of discontinuous embedded metal wires on the performance of an SRM with a single-burning-rate grain. A model based on convection heat transfer, heat conduction, and heat radiation was established to calculate the heat transfer in the discontinuous embedded metal wires in the grain, to then obtain the burning rate ratio. Most importantly, a solid rocket motor was designed to verify the feasibility of variable thrust and of the present model prediction, with the embedded silver–nickel alloy wire divided into two segments in the grain. According to the SRM ignition experiment, the silver–nickel alloy wires raised the burning rate of the grain. The pressure varied regularly with changes in the discontinuous embedded metal wires. The theoretical burning rate ratio matched the experimental result well. Based on the verified model, the effects of the burning rate, pressure exponent, burning rate ratio, and number of wires on thrust were investigated. Burning rate, burning rate ratio, and pressure exponent were found to be positively correlated with thrust ratio. The thrust ratio could reach 12.5 when the burning rate ratio was 5. The ability to adjust thrust tended to increase with an increase in the number of wires. This study also provided a method to assess whether the consecutive embedded metal wires had been broken or not. The method using discontinuous embedded metal wires in the grain was proven to be feasible to realize multi-thrusts of single-burning-rate grain, which is a new idea for the design of a multi-thrust SRM. Full article

(This article belongs to the Special Issue Combustion Evaluation and Control of Solid Rocket Motors)

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13 pages, 4418 KiB

Open AccessArticle

A Study on Ultra-Low-Pressure Ratio Technology on the Basis of 3D-Printed Propellant for a Solid Rocket Motor

by Shixiong Song, Quanbin Ren, Min Tang, Jiawei Shi and Jiawei Wang

Aerospace 2023, 10(10), 862; https://doi.org/10.3390/aerospace10100862 - 29 Sep 2023

Cited by 1 | Viewed by 1663

Abstract

Fused deposition technology (FDM), as an additive manufacturing (AM) technology, holds immense potential in the field of solid grain manufacturing. It can accomplish complex grain shaping with ultra-low-pressure ratios, which are challenging to achieve using conventional grain manufacturing processes. In this work, solid propellants with complex structures were made by using 3D printing. The obtained sample grains of the solid propellants had a complete structure, which conformed to the design model and had no obvious defects. Then, the combustion and mechanical properties of the printed solid propellant were obtained and analyzed. The results show that the composition of the printed solid propellant is more uniform and the performance is better than that of the conventional solid propellant. In addition, by conducting a motor experiment, it was verified that the 3D-printed grains with complex structures have the characteristic of an “ultra-low pressure ratio”. The comparative analysis revealed that the maximum working pressure was reduced by about 19.5%, the bearing load of the shell was reduced, and the mass of the shell and other bearing parts was reduced by 11.5%. The research in this paper shows that 3D-printed solid propellant technology can realize the formation of grains with complex structure, which can directly promote the solid rocket motor to obtain the “ultra-low pressure ratio” characteristic, and greatly improve the performance of solid rocket motors. Full article

(This article belongs to the Special Issue Combustion Evaluation and Control of Solid Rocket Motors)

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15 pages, 3392 KiB

Open AccessArticle

Using the Impulse Method to Determine High-Pressure Dynamic Burning Rate of Solid Propellants

by Jiahao Liu, Yinghong Wang, Xinyang Li and Junhao Cong

Aerospace 2023, 10(9), 818; https://doi.org/10.3390/aerospace10090818 - 18 Sep 2023

Cited by 2 | Viewed by 2267

Abstract

A new method for determining the burning rate of a solid propellant, called the Impulse Method, is proposed in this paper. It is based on the proportional relationship between the impulse generated and the mass of the burned propellant. The pressure–time and thrust–time curves are obtained from a tubular propellant grain burning in the chamber, whose inner surface serves as the initial burning surface. Consequently, the mass of the propellant that was burned off at different pressures can be determined, and the burning rates at different pressures are derived according to the geometric parameters of the propellant grain. The Impulse Method was applied to test the burning rate of two types of propellants twice. The results show that the burning rates were consistent for the same propellant at corresponding pressures, demonstrating the feasibility and reliability of the Impulse Method. The burning rate of a GAP-based composite propellant at 20 MPa measured using the Standard Motor Method was 22.6 mm/s, and that measured using the Impulse Method was 22.2 mm/s and 22.7 mm/s, respectively. These findings indicate that the two methods have comparable accuracy. However, the Impulse Method has the advantage of obtaining the burning rate of the solid propellant at any pressure through a single test. In addition, the nozzle erosion only affected the pressure and not the burning rate. Finally, the rationality of the approach for determining the actual specific impulse was proven by comparing the results with those from another testing method. Full article

(This article belongs to the Special Issue Combustion Evaluation and Control of Solid Rocket Motors)

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13 pages, 4793 KiB

Open AccessArticle

A Phenomenological Model for the Unsteady Combustion of Solid Propellants from a Zel’dovich-Novzhilov Approach

by Zhuopu Wang, Wenchao Zhang and Yuanzhe Liu

Aerospace 2023, 10(9), 767; https://doi.org/10.3390/aerospace10090767 - 29 Aug 2023

Cited by 2 | Viewed by 1150

Abstract

Solid rocket motors are prone to combustion instabilities, which may lead to various problems for the rockets, from unexpected oscillations, precision decreasing, to explosion. The unsteady combustion dynamics of the propellants play a crucial role in most solid rocket motors experiencing combustion instabilities. A modeling framework for the unsteady combustion of the solid propellant is constructed via the Zel’dovich-Novozhilov (ZN) phenomenological perspective. The overall unsteady combustion features of a quasi-steady homogeneous one-dimensional (QSHOD) model are investigated. The phenomenological ZN parameters are then calculated. Compared with the traditional ZN-QSHOD linear equivalence relation, the new calculated system yields better results for the pressure coupling response, especially in the non-linear regime. The proposed phenomenological modeling provides a new methodology for the model reduction of the complex flame models. Full article

(This article belongs to the Special Issue Combustion Evaluation and Control of Solid Rocket Motors)

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16 pages, 5936 KiB

Open AccessArticle

Numerical Investigation on the Effect of Ammonium Perchlorate Content and Position on the Combustion Characteristics of an Ammonium Perchlorate/Hydroxyl-Terminated Polybutadiene Propellant

by Di Sun, Yongzhou Li, Peijin Liu, Bofeng Chen and Wei Fan

Aerospace 2023, 10(8), 692; https://doi.org/10.3390/aerospace10080692 - 3 Aug 2023

Cited by 3 | Viewed by 1852

Abstract

A gas–solid-coupled sandwich combustion model was established for ammonium perchlorate (AP)/hydroxyl-terminated polybutadiene (HTPB) composite propellant. Numerical simulations were conducted to investigate the influence of the content of AP and the relative position of the coarse AP on the flame structure and the burning rate of the propellant. The results indicated that the overall AP mass fraction has a significant effect on the gas-phase flame temperature and burning rate, and there exists an optimal oxygen-to-fuel ratio that maximizes the burning rate. As the mass fraction of fine AP increased, the premixed flame above the binder matrix gradually took over the dominance of the diffusion flame, and the intensity of the diffusion flame near the interface of coarse AP and binder matrix also increased, resulting in a significant increase in the burning rate. As the mass fraction of fine AP increases from 0% to 70.0%, the average surface temperature increases from 937 K to 1026 K, and the burning rate rises from 0.9 cm/s to 2.7 cm/s. The location of the coarse AP causes the flame tilts to the side with less binder matrix, but it had little effect on the burn rate of the propellant. Full article

(This article belongs to the Special Issue Combustion Evaluation and Control of Solid Rocket Motors)

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12 pages, 3343 KiB

Open AccessArticle

Agglomerate Size Evolution in Solid Propellant Combustion under High Pressure

by Songchen Yue, Lu Liu, Huan Liu, Yanfeng Jiang, Peijin Liu, Aimin Pang, Guangxue Zhang and Wen Ao

Aerospace 2023, 10(6), 515; https://doi.org/10.3390/aerospace10060515 - 30 May 2023

Cited by 6 | Viewed by 2235

Abstract

Solid propellant combustion and flow are significantly affected by condensed combustion products (CCPs) in solid rocket motors. A new aluminum agglomeration model is established using the discrete element method, considering the burning rate and formulation of the propellant. Combining the aluminum combustion and alumina deposition model, an analytical model of the evolution of CCPs is proposed, capable of predicting the particle-size distribution of completely burned CCPs. The CCPs near and away from the propellant burning surface are collected by a special quench vessel under 6~10 MPa, to verify the applicability of the CCP evolution model. Experimental results show that the predicted error of the proposed CCP evolution model is less than 8.5%. Results are expected to help develop better analytical tools for the combustion of solid propellants and solid rocket motors. Full article

(This article belongs to the Special Issue Combustion Evaluation and Control of Solid Rocket Motors)

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19 pages, 7791 KiB

Open AccessArticle

Flow Coefficient and Starting Performance Prediction of Variable Geometry Curved Axisymmetric Inlet

by Yongzhou Li, Di Sun, Zejun Wu and Kunyuan Zhang

Aerospace 2023, 10(6), 506; https://doi.org/10.3390/aerospace10060506 - 27 May 2023

Cited by 2 | Viewed by 1538

Abstract

With the development of combined cycle engines, it is urgent to estimate more quickly and accurately the flow capture capacity and starting performance of variable geometry inlets over a wide Mach number range. Based on the flow field and parameter fitting, two prediction methods for the curved axisymmetric inlet with lip translation scheme have been proposed. The method based on the flow field of the reference inlet is more efficient than the parameters-based prediction method, as it can accurately predict the lip translation distance and the corresponding flow coefficient over the entire working range of the inlet without additional numerical calculations. Moreover, the starting Mach number is accurately predicted by the fitting method based on the throat Mach number of the reference inlet, with a relative error of only 0.95% compared to the numerical simulation. The flow coefficient-based method is simple and accurate for predicting lip translation distances with a known starting Mach number, with a relative error of only 1.65% compared to numerical simulations. The prediction approaches can overcome the drawbacks of the standard iterative algorithms and significantly enhance computational accuracy and efficiency. Full article

(This article belongs to the Special Issue Combustion Evaluation and Control of Solid Rocket Motors)

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20 pages, 18394 KiB

Open AccessArticle

Laser-Induced Ignition and Combustion of Single Micron-Sized Al-Li Alloy Particles in High Pressure Air/N₂

by Dunhui Xu, Fang Wang, Shengji Li, Xuefeng Huang, Heping Li and Yanhui Guo

Aerospace 2023, 10(3), 299; https://doi.org/10.3390/aerospace10030299 - 17 Mar 2023

Cited by 10 | Viewed by 2586

Abstract

To solve the problems associated with micron-sized aluminum (Al), including sintering, agglomeration, and slag deposition during the combustion of aluminized propellants, aluminum–lithium (Al-Li) alloy, prepared by introducing a small amount of Li (1.0 wt.%) into Al, was used in place of Al. Then, the ignition and combustion characteristics of single micron-sized Al-Li alloy particles were investigated in detail using a self-built experimental apparatus and multiple characterization methods. The ignition probability, ignition delay time, flame propagation rate, burn time, combustion temperature, flame radiation spectra, and microexplosion characteristics were obtained. The TG-DSC results demonstrated that, as compared to the counterpart Al, the Al-Li alloy had a lower ignition temperature. The emission lines of AlO revealed the gas-phase combustion of the Al-Li alloy, and thus the Al-Li alloy exhibited a mixed combustion mode, including surface combustion and gas-phase combustion. Moreover, during combustion, a microexplosion occurred, which increased the combustion rate and reduced the burn lifetime. The ambient pressure had a significant effect on the ignition and combustion characteristics of the Al-Li alloy, and the ignition delay time and burn time exponentially decreased as the ambient pressure enhanced. The combustion temperature of the Al-Li alloy at atmospheric pressure was slightly higher than those at elevated pressures. The Al-Li alloy burned in N₂, but no microexplosion was observed. Finally, the ignition and combustion mechanism of the Al-Li alloy in air was demonstrated by combining SEM, EDS, and XRD analyses of the material and residues. The results suggest that the addition of Li promoted the combustion performance of Al by changing the surface structure of the oxide film and the combustion mode. Full article

(This article belongs to the Special Issue Combustion Evaluation and Control of Solid Rocket Motors)

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12 pages, 1698 KiB

Open AccessArticle

Effect of Al–Li Alloy on the Combustion Performance of AP/RDX/Al/HTPB Propellant

by Weiqiang Xiong, Yunjie Liu, Tianfu Zhang, Shixi Wu, Dawen Zeng, Xiang Guo and Aimin Pang

Aerospace 2023, 10(3), 222; https://doi.org/10.3390/aerospace10030222 - 25 Feb 2023

Cited by 9 | Viewed by 2993

Abstract

Aluminium–lithium alloy (Al–Li alloy) powder has excellent ignition and combustion performance. The combustion product of Al–Li alloy powder combined with ammonium perchlorate is gaseous at the working temperature of solid rocket motors, which greatly reduces the loss of two-phase flow. Experimental investigations were thoroughly conducted to determine the effect of the Al–2.5Li (2.5 wt% lithium) content on propellant combustion and agglomeration based on thermogravimetry-differential scanning calorimetry, heat combustion, laser ignition, combustion diagnosis, a simulated 75 mm solid rocket motor and a condensed combustion products (CCPs) collection device. The results show that the exothermic heat and weight gain upon the thermal oxidation of Al–Li alloy is obviously higher than those of Al powder. Compared with the reference propellant’s formulation, Al–2.5Li leads to an increase in the burning rate and a decrease in the size of the condensed combustion products of the propellants. As the Al–2.5Li alloy content gradually increases from 0 wt% to 19 wt%, the burning rate increases from 5.391 ± 0.021 mm/s to 7.244 ± 0.052 mm/s at 7 MPa of pressure; meanwhile, the pressure exponent of the burning rate law is changed from 0.326 ± 0.047 to 0.483 ± 0.045, and the d₄₃ of the combustion residue is reduced from 165.31 ± 36.18 μm to 12.95 ± 4.00 μm. Compared to the reference propellant’s formulation, the combustion efficiency of the HTPB propellant is increased by about 4.4% when the Al–2.5Li alloy content is increased from 0 to 19%. Therefore, Al–2.5Li alloy powder is a promising fuel for solid propellants. Full article

(This article belongs to the Special Issue Combustion Evaluation and Control of Solid Rocket Motors)

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16 pages, 8223 KiB

Open AccessArticle

Study on Burning Surface Regression Algorithm under Erosive Burning Based on CT Images of Solid Rocket Motor Grain

by Shun Liu, Hongyi Lu, Bin Zhang, Yucheng Yang and Doudou Sang

Aerospace 2023, 10(1), 21; https://doi.org/10.3390/aerospace10010021 - 26 Dec 2022

Cited by 2 | Viewed by 3005

Abstract

The presence of the erosive burning effect during the operation of a solid rocket motor (SRM) is one of the most important factors affecting the proper operation of the motor. To solve the effects of the operating process of the motor under erosive burning, a synthesis algorithm based on the actual CT images is proposed to combine the Level-set (LS) method with the minimum distance function (MDF) method for the simulation of the burning surface regression of the grain under erosive burning. The Hamilton–Jacobi control equation can be solved exactly for the discrete form of LS. To improve the computational efficiency of the LS method, the minimum distance field is initialized and only the distortion grid is adjusted during the reinitialization. The third-order TVD Runge–Kutta method is used to solve the problem of numerical oscillation and improve the calculation accuracy. The experiments simulate the burning process of the NAWC No. 6 partial grain under erosive burning, which can provide the main basis for the performance design of solid propellant. The experimental results show that the method has good applicability to three-dimensional complex grains. It can realize the simulation of grains under erosive burning and its calculation accuracy is high. Full article

(This article belongs to the Special Issue Combustion Evaluation and Control of Solid Rocket Motors)

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18 pages, 5157 KiB

Open AccessArticle

Study on Separation Characteristics of Nozzles with Large Expansion Ratio of Solid Rocket Motors

by Zhihong Wang, Chunguang Wang and Weiping Tian

Aerospace 2023, 10(1), 4; https://doi.org/10.3390/aerospace10010004 - 21 Dec 2022

Cited by 2 | Viewed by 2667

Abstract

In order to study the flow characteristics of a nozzle with large expansion ratio and its influence on the force on the nozzle, ground cold flow test research and a fluid–structure coupling simulation analysis were carried out (maximum expansion ratio ε = 30.25). The variation and pulsation characteristics of the pressure near the measuring point area and the inlet pressure were obtained through experiments. Through the analysis of the peak-to-peak value and average value, it was found that the average pressure after separation increases by 90%, but the peak-to-peak value increases by about five times, indicating that the pressure fluctuation after separation is much larger than before separation. The separation flow field under cold flow conditions was simulated using the CFD commercial calculation software Fluent to verify the correctness of the numerical calculations. The fluid–structure coupling analysis was carried out on a large expansion ratio (maximum expansion ratio ε = 48) full-scale nozzle, and the structural deformation characteristics of the nozzle under the separation conditions were studied. The research results show that flow separation occurs in the nozzle with a large expansion ratio under ground conditions. Before the separation point, the pressure pulsation on the nozzle wall is small, and the turbulent pulsation effect is weak. After the separation point, the pressure pulsation increases, and the turbulent pulsation effect is enhanced. When the total pressure decreases, the separation area of the nozzle increases, and the separation flow field presents a strong asymmetry. Reducing the total inlet pressure by half resulted in approximately 50 times the lateral load. Under the combined influence of the ground conditions and low total pressure, the large lateral load caused by the asymmetry of the separation flow field will cause the deformation of the nozzle structure to increase by 5.5 times. This research provides an important reference for the design and experiment of nozzles with a large expansion ratio. Full article

(This article belongs to the Special Issue Combustion Evaluation and Control of Solid Rocket Motors)

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