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Journal articles on the topic 'Solid propellant rockets Combustion'

Author: Grafiati

Published: 4 June 2021

Last updated: 1 February 2022

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1

Kozin, V. S. "Effect of the thermal and gas-dynamic properties of solid rocket propellant particles on the propellant combustion rate." Technical mechanics 2021, no. 1 (April 30, 2021): 63–67. http://dx.doi.org/10.15407/itm2021.01.063.

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The aim of this work is to eliminate the explosion possibility of a rocket engine that operates on a fast-burning solid propellant. The problem is considered by analogy with experiments conducted earlier. Various ways to increase the propellant combustion rate are presented. Examples of how the solid propellant combustion rate depends on the metal fuel and the oxidizer particle size are given. It is shown that unstable combustion of a solid propellant at high combustion chamber pressures is due to unstable combustion of the gas phase in the vicinity of the bifurcation point. Zeldovich’s theory of nonstationary powder combustion is applied to analyzing the explosion dynamics of the Hrim-2 missile’s solid-propellant sustainer engine. This method of analysis has not been used before. The suggested version that this phenomenon is related to the aluminum particle size allows one to increase the combustion rate in the combustion chamber of a liquid-propellant engine, thus avoiding the vicinity of the bifurcation point. The combustion of solid propellants differing in aluminum particle size is considered. The metal fuel and the oxidizer particle sizes most optimal in terms of explosion elimination are determined and substantiated. The use of submicron aluminum enhances the evaporation of ammonium perchlorate due to the infrared radiation of aluminum particles heated to an appropriate radiation temperature. This increases the gas inflow into the charge channel, thus impeding the suppression of ammonium perchlorate sublimation by a high pressure, which is important in the case where the engine body materials cannot withstand a high pressure in the charge channel. This increases the stability and rate of solid propellant combustion. It is shown that the Hrim-2 missile’s solid propellant cannot be used in the Hran missile. The combustion rate is suggested to be increased by using fine-dispersed aluminum in the solid propellant.

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2

Pang, W. Q., F. Q. Zhao, L. T. DeLuca, C. Kappenstein, H. X. Xu, and X. Z. Fan. "Effects of Nano-Sized Al on the Combustion Performance of Fuel Rich Solid Rocket Propellants." Eurasian Chemico-Technological Journal 18, no. 3 (November 5, 2016): 197. http://dx.doi.org/10.18321/ectj425.

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Several industrial- and research – type fuel rich solid rocket propellants containing nano-metric aluminum metal particles, featuring the same nominal composition, were prepared and experimentally analyzed. The effects of nano-sized aluminum (nAl) on the rheological properties of metal/HTPB slurries and fuel rich solid propellant slurries were investigated. The energetic properties (heat of combustion and density) and the hazardous properties (impact sensitivity and friction sensitivity) of propellants prepared were analyzed and the properties mentioned above compared to those of a conventional aluminized (micro-Al, mAl) propellant. The strand burning rate and the associated combustion fl ame structure of propellants were also determined. The results show that nAl powder is nearly “round” or “ellipse” shaped, which is different from the tested micrometric Al used as a reference metal fuel. Two kinds of Al (nAl and mAl) powder can be dispersed in HTPB binder suffi ciently. The density of propellant decreases with increasing mass fraction of nAl powder; the measured heat of combustion, friction sensitivity, and impact sensitivity of propellants increase with increasing mass fraction of nAl powder in the formulation. The burning rates of fuel rich propellant increase with increasing pressure, and the burning rate of the propellant loaded with 20% mass fraction of nAl powder increases 77.2% at 1 MPa, the pressure exponent of propellant increase a little with increasing mass fraction of nAl powder in the explored pressure ranges.

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3

Aziz, Amir, Rizalman Mamat, Wan Khairuddin Wan Ali, and Mohd Rozi Mohd Perang. "Review on Typical Ingredients for Ammonium Perchlorate Based Solid Propellant." Applied Mechanics and Materials 773-774 (July 2015): 470–75. http://dx.doi.org/10.4028/www.scientific.net/amm.773-774.470.

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Ammonium perchlorate (AP) based solid propellant is a modern solid rocket propellant used in various applications. The combustion characteristics of AP based composite propellants were extensively studied by many research scholars to gain higher thrust. The amount of thrust and the thrust profile, which may be obtained from a specific grain design, is mainly determined by the propellant composition and the manufacturing process that produces the solid propellant. This article is intended to review and discuss several aspects of the composition and preparation of the solid rocket propellant. The analysis covers the main ingredients of AP based propellants such as the binder, oxidizer, metal fuel, and plasticizers. The main conclusions are derived from each of its components with specific methods of good manufacturing practices. In conclusion, the AP based solid propellant, like other composite propellants is highly influenced by its composition. However, the quality of the finished grain is mainly due to the manufacturing process.

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4

GARCÍA-SCHÄFER, J. E., and A. LIÑÁN. "Longitudinal acoustic instabilities in slender solid propellant rockets: linear analysis." Journal of Fluid Mechanics 437 (June 22, 2001): 229–54. http://dx.doi.org/10.1017/s0022112001004323.

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To describe the acoustic instabilities in the combustion chambers of laterally burning solid propellant rockets the interaction of the mean flow with the acoustic waves is analysed, using multiple scale techniques, for realistic cases in which the combustion chamber is slender and the nozzle area is small compared with the cross-sectional area of the chamber. Associated with the longitudinal acoustic oscillations we find vorticity and entropy waves, with a wavelength typically small compared with the radius of the chamber, penetrating deeply into the chamber. We obtain a set of differential equations to calculate the radial and axial dependence of the amplitude of these waves. The boundary conditions are provided by the acoustic admittance of the propellant surface, given by an existing analysis of the thin gas-phase reaction layer adjacent to the solid–gas interface, and of the nozzle, accounting here for the possible effect of the vorticity and entropy waves. The equations are integrated in closed form and the results provide the growth rate of the disturbances, which we use to determine the conditions for instability of the longitudinal oscillations.

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5

Sheikholeslam, Mohammad Reza Zadeh, Daryoosh Kazemi, and Hooman Amiri. "Experimental Analysis of the Influence of Length to Diameter Ratio on Erosive Burning in a Solid Tubular Propellant Grain." Applied Mechanics and Materials 110-116 (October 2011): 3394–99. http://dx.doi.org/10.4028/www.scientific.net/amm.110-116.3394.

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Erosive burning usually refers to the increase in the propellant burning rate caused by high velocity combustion gasses flowing over the propellant surface. It may seriously affect the performance of solid-propellant rocket motors [1]. A series of experiments had been made to study the effects of length to the diameter ratio in a single tubular propellant grain on the erosive burning phenomenon. In the same combustion pressure and different grain geometries, the burning pattern ofAP1based propellantwere recorded. Furthermore, pressure-time curve for each condition was obtained. The mean velocity gradient is obtained by some thermo-gas-dynamical analysis on experimental data. The results can be used for preliminary design ofAPbased tubular propellant rocket motors. This method may be used for other types of tubular solid propellants which defer in chemical formulation.

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6

Murachman, Bardi, Sajono Sajono, Fauzan Afandi, and Johan Khaeri. "Optimization Study of the Solid Propellant (Rocket Fuel) Based on Extracted Bitumen of Indonesian Natural Buton Asphalt." ASEAN Journal of Chemical Engineering 13, no. 2 (September 17, 2014): 57. http://dx.doi.org/10.22146/ajche.49732.

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The asphalt propellant for rockets has been investigated since 1960. This material has been developed with the variation of fuels, oxidizer, binders, metal elements and additives. As solid propellant, it has some advantages and disadvantages during the implementation. At present, Extracted Buton asphalt has been studied as an alternative propellant fuels. It is a natural asphalt, extracted from Buton island asphalt rock. When the extract of buton asphalt is mixed with oxidizer, binder, and metal powder, it can be functioned as propellant which is able to release high intensity of energy, have strong thrust and power to fly the rocket. This optimization study of solid propellant was conducted by mixing the Buton asphalt as fuel, oxidizer, metal element and other additives to form a solid propellant. The oxidizer consisted of potassium nitrate (KNO3) and potassium perchlorate (KClO4). The variations of KClO4/KNO3, propellant density and the ratio of the nozzle diameter were also conducted in order to find the best propellant composition and the optimum operating conditions to produce enough power while maintain the integrity of the rocket. The main parameters such as the propellant’s thrust (F) and the specific impulse (Isp) were examined. The results showed that higher composition of KClO4/KNO3 gave the higher value of the thrust and the specific impulse. KClO4/KNO3 levels above the 1:1 ratio produced an explosive properties at the time of ignition. The tendency of propellant to explode during ignition process was also observed. The optimum condition was obtained at the KClO4/KNO3 ratio of 1:1 , the propellant density was 1.900 gram/cm3 and Ae/A* was 3.33. These conditions generated impulse value that last for 277.07 seconds, average thrust of 14.082 N, and average rate of combustion of 0,24 cm/second. Therefore, it can be concluded that propellant with fuel from extracted of Buton asphalt can be used as an alternative propellant for rocket.

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7

Oyedeko K.F.K and Egwenu S. O. "Modelling of the formulated solid rocket propellant characteristics." Global Journal of Engineering and Technology Advances 6, no. 2 (February 28, 2021): 061–73. http://dx.doi.org/10.30574/gjeta.2021.6.2.0017.

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This study is a mathematical model to obtain the characteristics performance of magnesium metal (powder) and carbon on a potassium nitrate-sucrose (KNSU) solid propellant formulation. Characterization of propellant is, as a general rule, important to determine its performance before it can be suitable for use for a rocket flight or any mission. Method of ballistic load cell evaluation was used to validate results and a mathematical model using the combustion exhaust products was solved to obtain the characteristics performance parameters of the propellant. The carbon constituent which acts as an opacifier and coolant was kept constant at 2% in order to arrest some of the heat during the combustion process and helped to lower the combustion temperature, because high combustion temperature could lead to combustion chamber rupture or failure. The effect of addition of magnesium which was optimized for 3% in the formulation contributed significantly in improving the overall performance of the propellant. The utilization of magnesium in KNSU propellant provided higher values parameters and better performance compared to when not included. This was confirmed with the model equations. The propellant combustion products equation was used to model and obtain the characteristics performance parameters. This gave propellant specific impulse (122.9s), combustion temperature (1821K), heat ratio (1.1592), molecular weight (36.89g/mole), propellant density (1912.5kg/m3) and characteristics velocity (1000m/s) result while maintaining the same chamber pressure.

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8

Oyedeko K.F.K and Egwenu S. O. "Effect of magnesium metal in the characteristics performance of a sucrose-based solid rocket propellant." Global Journal of Engineering and Technology Advances 6, no. 2 (February 28, 2021): 051–60. http://dx.doi.org/10.30574/gjeta.2021.6.2.0016.

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This research work aimed at investigating the effects of magnesium metal (powder) and carbon on a potassium nitrate-sucrose (KNSU) solid propellant formulation. Characterization of propellant is very important to determine its performance before it can be suitable for use for a rocket flight or any mission. Ballistic loadcell method was used. The ballistic load cell instrumentation measured the thrust generated by the propellant, the propellant burn time and the exit temperature of the burning hot propellant gases. The carbon constituent which acts as an opacifier and coolant was kept constant at 2% in order to arrest some of the heat during the combustion process and helped to lower the combustion temperature, because high combustion temperature could lead to combustion chamber rupture or failure. Also, carbon was not increased beyond 2%, so as not to make the propellant excessively smoky because of presence of magnesium oxide and other solids in the combustion products that can cause air pollution, and could be harmful to human lives and the environment. The propellant specific impulse (117.9s), combustion temperature (1818K), heat ratio (1.1508), propellant molecular weight (38.88g/mole), propellant density (1874.6kg/m3), characteristics velocity (997.2m/s) and burn rate (0.00906m/s) were obtained. The effect of addition of magnesium which was optimized for 3% in the formulation contributed significantly in improving the overall performance of the propellant as parameters such as the specific impulse, chamber temperature, characteristics velocity and heat ratio were found to have higher values as compare to the KNSU propellant when magnesium was not present in the formulation. Basically, higher values of these parameters suggest better propellant performance. Also, in this case, when carbon was increased beyond 2%, the propellant was excessively smoky because of presence of magnesium oxide and other solids in the combustion products that can cause air pollution, and could be harmful to human lives and the environment.

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9

Bogusz, Rafał, Paulina Magnuszewska, and Bogdan Florczak. "STUDIES OF HIGH EXPLOSIVES IMPACT ON REDUCTION OF HCL IN HETEROGENEOUS SOLID ROCKET PROPELLANTS." PROBLEMY TECHNIKI UZBROJENIA, no. 3 (December 1, 2017): 29–45. http://dx.doi.org/10.5604/01.3001.0010.6308.

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The paper describes an influence of high explosives: hexogene (RDX), octogene (HMX), and dinitro-diaminoethene (FOX-7) on the properties of heterogeneous solid rocket propellant (HSRP) prepared on the base of Hydroxy Terminated Polybutadiene (HTPB) in which ammonium perchlorate (AP) was partially replaced by sodium nitrate (SN). It reduced the content of HCl in combustion products. Theoretical values of thermochemical and thermodynamic properties such as isochoric combustion heat (Q), specific impulse (Isp) and contents of combustion products in motor combustion chamber and nozzle have been identified by using the ICT-Code program. The rheological properties (virtual viscosity) of the propellant slurry during curing process, the sensitivity to mechanical stimuli (impact, friction), decomposition temperature, calorific value and hardness of propellants containing explosive materials were tested by instruments and ballistic properties were investigated by laboratory rocket motor (LRM).

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10

Kurdyumov, Vadim N. "Steady Flows in the Slender, Noncircular, Combustion Chambers of Solid Propellant Rockets." AIAA Journal 44, no. 12 (December 2006): 2979–86. http://dx.doi.org/10.2514/1.21125.

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11

Korotkikh, Alexander, Ivan Sorokin, and Ekaterina Selikhova. "Ignition and combustion of high-energy materials containing aluminum, boron and aluminum diboride." MATEC Web of Conferences 194 (2018): 01055. http://dx.doi.org/10.1051/matecconf/201819401055.

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Boron and its compounds are among the most promising metal fuel components to be used in solid propellants for solid fuel rocket engine and ramjet engine. Papers studying boron oxidation mostly focus on two areas: oxidation of single particles and powders of boron, as well as boron-containing composite solid propellants. This paper presents the results of an experimental study of the ignition and combustion of the high-energy material samples based on ammonium perchlorate, ammonium nitrate, and an energetic combustible binder. Powders of aluminum, amorphous boron and aluminum diboride, obtained by the SHS method, were used as the metallic fuels. It was found that the use of aluminum diboride in the solid propellant composition makes it possible to reduce the ignition delay time by 1.7–2.2 times and significantly increase the burning rate of the sample (by 4.8 times) as compared to the solid propellant containing aluminum powder. The use of amorphous boron in the solid propellant composition leads to a decrease in the ignition delay time of the sample by a factor of 2.2–2.8 due to high chemical activity and a difference in the oxidation mechanism of boron particles. The burning rate of this sample does not increase significantly.

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12

Novozhilov, B. V. "NONLINEAR COMBUSTION IN SOLID PROPELLANT ROCKET MOTORS." International Journal of Energetic Materials and Chemical Propulsion 5, no. 1-6 (2002): 793–802. http://dx.doi.org/10.1615/intjenergeticmaterialschemprop.v5.i1-6.830.

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13

Glebov, G. A., and S. A. Vysotskaya. "On the question of solid-propellant rocket engine design preventing unstable operation in the combustion chamber." Journal of «Almaz – Antey» Air and Space Defence Corporation, no. 4 (December 30, 2017): 63–72. http://dx.doi.org/10.38013/2542-0542-2017-4-63-72.

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The paper presents results of a numerical investigation concerning the effect that the flow duct shape and combustion rate equation have on the gas dynamic vortex flow pattern and self-excited pressure oscillations in the combustion chamber of a solid-propellant rocket engine. We provide guidelines on upgrading solid-propellant rocket engines in order to decrease the magnitude of pressure pulses in the case of pulsating combustion.

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14

Li, Chaolong, Zhixun Xia, Likun Ma, Xiang Zhao, and Binbin Chen. "Numerical Study on the Solid Fuel Rocket Scramjet Combustor with Cavity." Energies 12, no. 7 (March 31, 2019): 1235. http://dx.doi.org/10.3390/en12071235.

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Scramjet based on solid propellant is a good supplement for the power device of future hypersonic vehicles. A new scramjet combustor configuration using solid fuel, namely, the solid fuel rocket scramjet (SFRSCRJ) combustor is proposed. The numerical study was conducted to simulate a flight environment of Mach 6 at a 25 km altitude. Three-dimensional Reynolds-averaged Navier–Stokes equations coupled with shear stress transport (SST) k − ω turbulence model are used to analyze the effects of the cavity and its position on the combustor. The feasibility of the SFRSCRJ combustor with cavity is demonstrated based on the validation of the numerical method. Results show that the scramjet combustor configuration with a backward-facing step can resist high pressure generated by the combustion in the supersonic combustor. The total combustion efficiency of the SFRSCRJ combustor mainly depends on the combustion of particles in the fuel-rich gas. A proper combustion organization can promote particle combustion and improve the total combustion efficiency. Among the four configurations considered, the combustion efficiency of the mid-cavity configuration is the highest, up to about 70%. Therefore, the cavity can effectively increase the combustion efficiency of the SFRSCRJ combustor.

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15

KOSTYUSHIN, Kirill V. "NUMERICAL INVESTIGATION OF UNSTEADY GASDYNAMIC PROCESSES AT THE LAUNCH OF SOLID-PROPELLANT ROCKETS." Vestnik Tomskogo gosudarstvennogo universiteta. Matematika i mekhanika, no. 67 (2020): 127–43. http://dx.doi.org/10.17223/19988621/67/12.

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The paper presents the results of the methodology developed for calculating unsteady gasdynamic processes occurring at the launch of missiles, in the gas-dynamic paths of rocket engines, and in the external regions. The method accounts for the variation in the geometry of the solidpropellant charge in the course of solid-propellant rocket engine operation and in the geometry of the computational domain at the rocket launch. The analysis of the unsteady force impact of the supersonic jet on the launch surface is carried out. It is shown that the maximum force action is located in the vicinity of the Mach disks of the unperturbed jet. Numerical studies of gasdynamic processes at the launch of a model solid-propellant booster rocket are implemented including the case when the nozzle plug opening is taken into account. The contribution of the thrust force components at the stage of bootstrap operation is assessed. The presence of the plug at the initial stage of the engine start leads to an abrupt change in the thrust and minor fluctuations, which are damped as the pressure in the combustion chamber rises.

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16

Li, Jun-Qiang, Linlin Liu, Xiaolong Fu, Deyun Tang, Yin Wang, Songqi Hu, and Qi-Long Yan. "Transformation of Combustion Nanocatalysts inside Solid Rocket Motor under Various Pressures." Nanomaterials 9, no. 3 (March 6, 2019): 381. http://dx.doi.org/10.3390/nano9030381.

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In this paper, the dependences of the morphology, particle sizes, and compositions of the condensed combustion products (CCP) of modified double-base propellants (1,3,5-trimethylenetrinitramine (RDX) as oxidizer) on the chamber pressure (<35 MPa) and nickel inclusion have been evaluated under a practical rocket motor operation. It has been shown that higher pressure results in smaller average particle sizes of the CCPs. The CCPs of Ni-containing propellants have more diverse morphologies, including spherical particles, large layered structures, and small flakes coated on large particles depending on the pressure. The specific surface area (SSA) of CCPs is in the range of 2.49 to 3.24 m2 g−1 for propellants without nickel are less dependent on the pressure, whereas it is 1.22 to 3.81 Ni-based propellants. The C, N, O, Al, Cu, Pb, and Si are the major elements presented on the surfaces of the CCP particles of both propellants. The compositions of CCPs from Ni-propellant are much more diverse than another one, but only three or four major phases have been found for both propellants under any pressure. The metallic copper is presented in CCPs for both propellants when the chamber pressure is low. The lead salt as the catalyst has been transformed in to Pb(OH)Cl as the most common products of lead-based catalysts with pressure lower than 15 MPa. When pressure is higher than 5 MPa, the nickel-based CCPs has been found to contain one of the following crystalline phases: Pb2Ni(NO2)6, (NH4)2Ni(SO4)2·6H2O, C2H2NiO4·2H2O, and NiO, depending on the pressure.

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17

Xiao, Yumin, R. S. Amano, Timin Cai, and Jiang Li. "New Method to Determine the Velocities of Particles on a Solid Propellant Surface in a Solid Rocket Motor." Journal of Heat Transfer 127, no. 9 (April 19, 2005): 1057–61. http://dx.doi.org/10.1115/1.1999652.

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Use of aluminized composite solid propellants and submerged nozzles are common in solid rocket motors (SRM). Due to the generation of slag, which injects into a combusted gas flow, a two-phase flow pattern is one of the main flow characteristics that need to be investigated in SRM. Validation of two-phase flow modeling in a solid rocket motor combustion chamber is the focus of this research. The particles’ boundary conditions constrain their trajectories, which affect both the two-phase flow calculations, and the evaluation of the slag accumulation. A harsh operation environment in the SRM with high temperatures and high pressure makes the measurement of the internal flow field quite difficult. The open literature includes only a few sets of experimental data that can be used to validate theoretical analyses and numerical calculations for the two-phase flow in a SRM. Therefore, mathematical models that calculate the trajectories of particles may reach different conclusions mainly because of the boundary conditions. A new method to determine the particle velocities on the solid propellant surface is developed in this study, which is based on the x-ray real-time radiography (RTR) technique, and is coupled with the two-phase flow numerical simulation. Other methods imitate the particle ejection from the propellant surface. The RTR high-speed motion analyzer measures the trajectory of the metal particles in a combustion chamber. An image processing software was developed for tracing a slug particle path with the RTR images in the combustion chamber, by which the trajectories of particles were successfully obtained.

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18

Glebov, G. A., and S. A. Vysotskaya. "On the influence of the charge channel geometry and fuel properties on the working process instability in the solid propellant rocket combustion chamber." Journal of «Almaz – Antey» Air and Space Defence Corporation, no. 1 (March 30, 2017): 67–75. http://dx.doi.org/10.38013/2542-0542-2017-1-67-75.

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The paper numerically studies the influence of how the charge channel shape and the law of solid fuel burning rate influence the gas-dynamic vortex flow pattern and pressure self-oscillations in the combustion chamber of a solid-propellant rocket engine. The work presents the findings of the research which show that, using the numerical method, it is possible to choose the optimal shape of the charge channel and the solid propellant grade, which provide the least probability of occurrence of the pulsating combustion regime

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19

Abdelaziz, Almostafa, Liang Guozhu, and Anwer Elsayed. "Parameters Affecting the Erosive Burning of Solid Rocket Motor." MATEC Web of Conferences 153 (2018): 03001. http://dx.doi.org/10.1051/matecconf/201815303001.

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Increasing the velocity of gases inside solid rocket motors with low port-to-throat area ratios, leading to increased occurrence and severity of burning rate augmentation due to flow of propellant products across burning propellant surfaces (erosive burning), erosive burning of high energy composite propellant was investigated to supply rocket motor design criteria and to supplement knowledge of combustion phenomena, pressure, burning rate and high velocity of gases all of these are parameters affect on erosive burning. Investigate the phenomena of the erosive burning by using the 2’inch rocket motor and modified one. Different tests applied to fulfil all the parameters that calculated out from the experiments and by studying the pressure time curve and erosive burning phenomena.

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20

BABUK, V. A., V. A. VASSILIEV, and V. V. SVIRIDOV. "Propellant Formulation Factors and Metal Agglomeration in Combustion of Aluminized Solid Rocket Propellant." Combustion Science and Technology 163, no. 1 (February 2001): 261–89. http://dx.doi.org/10.1080/00102200108952159.

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21

Wang, Wei, Jiang Li, Ke Zhang, and Yang Liu. "Computing Method Investigation and Verification of Gas-Solid Combustion in Magnesium-Aluminum Based Propellant Ducted Rocket." Advanced Materials Research 503-504 (April 2012): 490–93. http://dx.doi.org/10.4028/www.scientific.net/amr.503-504.490.

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The combustion mechanism, which consisting of 22 species and 23 reaction equations, and three discrete models such as inertia, combusting, modification droplet, are employed for the investigation of gas-solid combustion in magnesium-aluminum based propellant ducted rocket based on thermal performance calculation. And path lines, temperature distribution, sediments are discussed after the computing method is validated by direct-connect experimentation and the flow field information, which obtained by numerical method and coincided with currently conclusions. The results indicated that the proposed method is reliable and practicable.

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22

Kim, Youngin, and Jeongho Cho. "Surface Erosion Analysis for Thermal Insulation Materials of Graphite and Carbon–Carbon Composite." Applied Sciences 9, no. 16 (August 13, 2019): 3323. http://dx.doi.org/10.3390/app9163323.

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A rocket uses fuel and oxidizers to generate propulsion by combustion and ejection, and is used for space exploration aircrafts, weapons, and satellite launches. In particular, the nozzle generating thrust of solid-propellant rockets is exposed to a high-temperature and high-pressure environment with erosion occurring from the combustion gas. When erosion occurs on the nozzle throat of such a rocket, it has a great impact on the flight performance such as reaching distance and flight speed. Many studies have been conducted to characterize erosion based on the thermochemical erosion model, since it has become important to choose nozzle materials suitable for such environments having robustness against combustion gasses of high temperature and high pressure. However, there is a limit to fully analyze the erosion characteristics only by the thermochemical erosion model. In this paper, we thus consider the mechanical erosion model with the thermochemical model for better understanding of erosion characteristics and investigate the thermochemical and mechanical erosion characteristics of nozzle throat heat-resistant materials made of graphite and carbon–carbon composites; the main factors affecting erosion are discussed by comparing the results of the experimental and theoretical models.

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23

Ma, Yanjie, 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." International Journal of Aerospace Engineering 2020 (December 8, 2020): 1–9. http://dx.doi.org/10.1155/2020/8889333.

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Erosive burning refers to the augmentation of propellant burning rate appears when the velocity of combustion gas flowing parallel to the propellant surface is relatively high. Erosive burning can influence the total burning rate of propellant and performance of solid rocket motors dramatically. There have been many different models to evaluate erosive burning rate for now. Yet, due to the complication processes involving in propellant and solid rocket motor combustion, unknown constants often exist in these models. To use these models, trial-and-error procedure must be implemented to determine the unknown constants firstly. This makes many models difficult to estimate erosive burning before plenty of experiments. In this paper, a new erosive burning rate model is proposed based on the assumption that the erosive burning rate is proportional to the heat flux at the propellant surface. With entrance effect, roughness, and transpiration considered, convective heat transfer coefficient correlation proposed in recent years is used to compute the heat flux. This allows the release of unknown constants, making the model universal and easy to implement. The computational data of the model are compared with different experimental and computational data from different models. Results show that good accuracy (10%) with experiments can be achieved by this model. It is concluded that the present model could be used universally for erosive burning rate evaluation of propellant and performance prediction of solid rocket motor as well.

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24

Larson, Richard S. "Prediction of aluminum combustion efficiency in solid propellant rocket motors." AIAA Journal 25, no. 1 (January 1987): 82–91. http://dx.doi.org/10.2514/3.9585.

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25

Kim, Hakchul, Junseong Kim, Heejang Moon, Honggye Sung, Hunki Lee, Wonsuk Ohm, and Dohyung Lee. "Linear Stability Analysis for Combustion Instability in Solid Propellant Rocket." Journal of the Korean Society of Propulsion Engineers 17, no. 5 (October 1, 2013): 27–36. http://dx.doi.org/10.6108/kspe.2013.17.5.027.

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26

Volpi, Angelo, and C. Zanotti. "INVESTIGATION OF SOLID ROCKET PROPELLANT COMBUSTION BY LASER DOPPLER ANEMOMETRY." International Journal of Energetic Materials and Chemical Propulsion 3, no. 1-6 (1994): 654–58. http://dx.doi.org/10.1615/intjenergeticmaterialschemprop.v3.i1-6.630.

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27

Lee, Hunki, Taeyoung Park, Won-Suk Ohm, and Dohyung Lee. "Nonlinear acoustics of combustion instability in solid-propellant rocket motors." Journal of the Acoustical Society of America 134, no. 5 (November 2013): 4100. http://dx.doi.org/10.1121/1.4830977.

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28

Novozhilov, B. V., Z. I. Kaganova, and A. A. Belyaev. "Simulation of unsteady combustion in a solid-propellant rocket motor." Russian Journal of Physical Chemistry B 3, no. 1 (February 2009): 91–98. http://dx.doi.org/10.1134/s1990793109010151.

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29

Yun, Myeong-Won, and Gyeong-Mu Kim. "Case Study of Combustion Instability in Solid Propellant Rocket Motors." Journal of the Korean Society for Aeronautical & Space Sciences 31, no. 1 (February 1, 2003): 133–40. http://dx.doi.org/10.5139/jksas.2003.31.1.133.

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30

Saha, S., and D. Chakraborty. "Computational Fluid Dynamics Simulation of Combustion Instability in Solid Rocket Motor : Implementation of Pressure Coupled Response Function." Defence Science Journal 66, no. 3 (April 25, 2016): 216. http://dx.doi.org/10.14429/dsj.66.9058.

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<p class="NomenclatureClauseTitle">Combustion instability in solid propellant rocket motor is numerically simulated by implementing propellant response function with quasi steady homogeneous one dimensional formulation. The convolution integral of propellant response with pressure history is implemented through a user defined function in commercial computational fluid dynamics software. The methodology is validated against literature reported motor test and other simulation results. Computed amplitude of pressure fluctuations compare closely with the literarture data. The growth rate of pressure oscillations of a cylindrical grain solid rocket motor is determined for different response functions at the fundamental longitudinal frequency. It is observed that for response function more than a critical value, the motor exhibits exponential growth rate of pressure oscillations.</p>

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31

Abrukov, Victor S., Alexander N. Lukin, Darya A. Anufrieva, Charlie Oommen, V. R. Sanalkumar, Nichith Chandrasekaran, and Rajaghatta Sundararam Bharath. "Recent Advancements in Study of Effects of Nano Micro Additives on Solid Propellants Combustion by Means of the Data Science Methods." Defence Science Journal 69, no. 1 (January 10, 2019): 20–26. http://dx.doi.org/10.14429/dsj.69.12948.

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The efforts of Russian-Indian research team for application of the data science methods, in particular, artificial neural networks for development of the multi-factor computational models for studying effects of additive’s properties on the solid rocket propellants combustion are presented. The possibilities of the artificial neural networks (ANN) application in the generalisation of the connections between the variables of combustion experiments as well as in forecasting of “new experimental results” are demonstrated. The effect of particle size of catalyst, oxidizer surface area and kinetic parameters like activation energy and heat release on the final ballistic property of AP-HTPB based propellant composition has been modelled using ANN methods. The validated ANN models can predict many unexplored regimes, like pressures, particle sizes of oxidiser, for which experimental data are not available. Some of the regularly measured kinetic parameters extracted from non-combustion conditions could be related to properties at combustion conditions. Results predicted are within desirable limits accepted in combustion conditions.

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32

Xianggeng, Wei, Bo Tao, Wang Pengbo, Ma Xinjian, Lou Yongchun, and Chen Jian. "Burning Rate Enhancement Analysis of End-Burning Solid Propellant Grains Based on X-Ray Real-Time Radiography." International Journal of Aerospace Engineering 2020 (June 22, 2020): 1–9. http://dx.doi.org/10.1155/2020/7906804.

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Unexpected pressure rise may occur in the end-burning grain solid rocket motor. It is generally believed that this phenomenon is caused by the nonparallel layer combustion of the burning surface, resulting in the increase of burning rate along the inhibitor. In order to explain the cause of this phenomenon, the experimental investigation on four different end configurations were carried out. Based on the X-ray real-time radiography (RTR) technique, a new method for determining the dynamic burning rate of propellant and obtaining the real-time end-burning profile was developed. From the real-time images of the burning surface, it is found that there was a phenomenon of nonuniform burning surface displacement in the end-burning grain solid rocket motor. Through image processing, the real-time burning rate of grain center line and the real-time cone angle are obtained. Based on the analysis of the real-time burning rate at different positions of the end surface, the end face cone burning process in the motor working process is obtained. The closer to the shell, the higher the burning rate of the propellant. Considering the actual structure of this end-burning grain motor, it is speculated that the main cause of the cone burning of the grain may be due to the heat conduction of the metal wall. By adjusting the initial shape of the grain end surface, the operating pressure of the combustion chamber can be basically unchanged, so as to meet the mission requirements. The results show that the method can measure the burning rate of solid propellant in real time and provide support for the study of nonuniform combustion of solid propellant.

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33

Kiryushkin, A. E., and L. L. Minkov. "NUMERICAL SIMULATION OF INTRA-CHAMBER PROCESSES IN A SOLID ROCKET MOTOR WITH ACCOUNT FOR BURNING SURFACE MOTION." Vestnik Tomskogo gosudarstvennogo universiteta. Matematika i mekhanika, no. 71 (2021): 90–105. http://dx.doi.org/10.17223/19988621/71/8.

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The axisymmetric solid rocket motor (SRM) with an “umbrella” shape is considered in this paper. The numerical algorithm based on the inverse Lax-Wendroff procedure for a gas dynamic equation and on the level-set method for tracking the burning surface is overviewed for internal ballistics problems. Assuming that the propellant combustion proceeds in a quasi-stationary regime and a mass flow from the burning surface depends on the pressure raised to the power of parameter ν, the numerical computations of intra-chamber combustion product flows during the main-firing phase are carried out using the numerical algorithm developed for “umbrella”-shaped SRM at different parameter values. The approximation convergence of flow parameters in a case of the stationary propellant surface and average intra-chamber pressure for all the time of motor operation is examined. The numerical simulation results are obtained and analyzed for different “umbrella” inclination angles. Though the developed algorithm has been applied to the motors with a specific shape, it can also be used for propellant grains of different shapes and is easily extended to 3D models.

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34

Hegab, A. M., H. H. Sait, A. Hussain, and A. S. Said. "Numerical modeling for the combustion of simulated solid rocket motor propellant." Computers & Fluids 89 (January 2014): 29–37. http://dx.doi.org/10.1016/j.compfluid.2013.10.029.

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35

Petrenko, V. I., and V. L. Popov. "A variable-thrust solid-propellant rocket engine with local combustion boosting." Combustion, Explosion, and Shock Waves 32, no. 3 (May 1996): 327–30. http://dx.doi.org/10.1007/bf01998464.

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36

Yoon, Jae-Kun. "Prediction of longitudinal combustion instability in a solid-propellant rocket motor." KSME Journal 8, no. 2 (June 1994): 206–13. http://dx.doi.org/10.1007/bf02953270.

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37

Gottlieb, J. J., and D. R. Greatrix. "Numerical Study of the Effects of Longitudinal Acceleration on Solid Rocket Motor Internal Ballistics." Journal of Fluids Engineering 114, no. 3 (September 1, 1992): 404–10. http://dx.doi.org/10.1115/1.2910045.

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The internal ballistics of a solid-propellant rocket motor subjected to both constant and oscillatory longitudinal accelerations are studied. The one-dimensional time-dependent equations of motion governing the unsteady two-phase core flow in the accelerating motor chamber and nozzle are solved numerically by using the random-choice method, along with pressure-dependent and crossflow-dependent burning-rate equations for propellant combustion. A constant forward acceleration produces negligible effects, whereas longitudinal motor vibrations near the natural frequency of waves criss-crossing the length of the motor chamber can produce large but bounded oscillatory motor-chamber pressures.

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38

Trushlyakov, V., K. Zharikov, and D. Lempert. "Development of Solid Gas Generating Compositions to Ensure Non Explosiveness of Spent Orbital Stages of Liquid Rocket of Space Launch Vehicles." Eurasian Chemico-Technological Journal 19, no. 1 (June 19, 2017): 63. http://dx.doi.org/10.18321/ectj606.

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The choice is discussed of solid gas generating compositions for venting by hot combustion products a fuel tank of the spent orbital stage of a space launch vehicle with a main liquid rocket engine. Non explosiveness is achieved via eliminating the<br />possibility of freezing the drainage system when products of gasification (vapours of a propellant component + the remains of a gas boost + the hot products of combustion of solid gas generating compositions) are discharged from the tank into surrounding space. There are imposed requirements, constraints, and criteria for selecting solid gas generating compositions. When considering tank with the residues of liquid oxygen belonging to orbital spent stage of the launch vehicle «Zenith» the ways are shown how to ensure explosion safety, which on the basis of proposed approaches by selecting solid gas generating compositions (SGC) which generate oxygen and<br />nitrogen. As a criterion of choice of SGC the total mass of the gasification system is adopted, which includes the SGC mass for gasification of liquid propellant residues, the mass of the gas generator and the mass of system to supply the combustion products of SGC into the tank. It is proposed use of residual heat in the condensed phase of the SGC combustion products to heat up the drainage system, which will increase the probability of a trouble-free operation of the drainage system.

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39

Han, Qi Long, Zuo Ming Zhu, Xin Gao, and Qi Yun Zhang. "Run Risk Analysis of Solid Rocket Motor Waterjet Clearing System." Applied Mechanics and Materials 483 (December 2013): 660–63. http://dx.doi.org/10.4028/www.scientific.net/amm.483.660.

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According to the hazard identification theory, run risks of solid rocket motor waterjet clearing system were identified from three different angles and potential causes of run risks were investigated; based on analysis of accidents that may occur in the running of solid rocket motor waterjet clearing system, fault tree analysis of propellant combustion/explosion and AP wastewater spill was carried out, minimal cut sets (path sets) and structural importance of each basic event were obtained, and furthermore corresponding recommendations in the accident prevention of solid rocket motor waterjet clearing system running were given accordingly.

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40

Sabnis, Jayant S. "Numerical Simulation of Distributed Combustion in Solid Rocket Motors with Metalized Propellant." Journal of Propulsion and Power 19, no. 1 (January 2003): 48–55. http://dx.doi.org/10.2514/2.6101.

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41

Babuk, V. A., and V. A. Vasilyev. "Model of Aluminum Agglomerate Evolution in Combustion Products of Solid Rocket Propellant." Journal of Propulsion and Power 18, no. 4 (July 2002): 814–23. http://dx.doi.org/10.2514/2.6005.

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42

Li, Jiang, Kai Liu, Yuanhao Gao, Shichang Liu, Wei Wang, and Yang Liu. "Combustion Characteristics Experimental Study of Solid Hydrocarbon Propellant for Air-Turbo Rocket." Journal of Propulsion and Power 34, no. 5 (September 2018): 1198–205. http://dx.doi.org/10.2514/1.b36917.

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43

MORITA, Takakazu. "A concept for laser controlled combustion in a solid propellant rocket motor." Proceedings of Conference of Kanto Branch 2003.9 (2003): 419–20. http://dx.doi.org/10.1299/jsmekanto.2003.9.419.

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44

Rashkovskii, S. A., Yu M. Milekhin, A. V. Fedorychev, and I. G. Assovskii. "Stability of combustion in solid-propellant rocket motors with pressure stabilization system." Doklady Physical Chemistry 428, no. 1 (September 2009): 178–82. http://dx.doi.org/10.1134/s001250160909005x.

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45

Griego, Castillo, Nadir Yilmaz, and Alpaslan Atmanli. "Analysis of aluminum particle combustion in a downward burning solid rocket propellant." Fuel 237 (February 2019): 405–12. http://dx.doi.org/10.1016/j.fuel.2018.10.016.

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46

Rezaiguia, Hichem, Peijin Liu, and Tianhao Yang. "Flame response of solid propellant AP/Al/HTPB to a longitudinal acoustic wave." International Journal of Spray and Combustion Dynamics 9, no. 4 (May 3, 2017): 241–59. http://dx.doi.org/10.1177/1756827717695830.

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This paper is devoted to an experimental work which consists of the analysis of the flame of a small solid propellant sample, AP/Al/HTPB, subjected to a longitudinal acoustic wave. Experiments were conducted in a closed tube under two mean pressures: 1 and 2.5 MPa. The qualitative and quantitative analysis of the flame snapshots, using a microscope and a high-speed camera, revealed that the acoustic wave created at the end of the chamber by a pulser system strongly affects the flame and the combustion products dynamic above the solid propellant surface, namely, the flame and the hot products oscillate around a line perpendicular to the propellant surface. This dynamic of the hot gas disturbs the local burning rate and the regression surface profile. Thus, the thrust and the burning duration will change, therefore, the flight path of the rocket may shift and can lead to failure of the mission.

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47

Amado, Javier Carlos Quagliano, Pablo Germán Ross, Natália Beck Sanches, Juliano Ribeiro Aguiar Pinto, and Jorge Carlos Narciso Dutra. "Evaluation of elastomeric heat shielding materials as insulators for solid propellant rocket motors: A short review." Open Chemistry 18, no. 1 (December 5, 2020): 1452–67. http://dx.doi.org/10.1515/chem-2020-0182.

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AbstractThis review addresses a comparison, based on the literature, among nitrile rubber (NBR), ethylene-propylene-diene-monomer rubber (EPDM), and polyurethane (PU) elastomeric heat shielding materials (EHSM). Currently, these are utilized for the insulation of rocket engines to prevent catastrophic breakdown if combustion gases from propellant reaches the motor case. The objective of this review is to evaluate the performance of PU–EHSM, NBR–EHSM, and EPDM–EHSM as insulators, the latter being the current state of the art in solid rocket motor (SRM) internal insulation. From our review, PU–EHSM emerged as an alternative to EPDM–EHSM because of their easier processability and compatibility with composite propellant. With the appropriate reinforcement and concentration in the rubber, they could replace EPDM in certain applications such as rocket motors filled with composite propellant. A critical assessment and future trends are included. Rubber composites novelties as EHSM employs specialty fillers, such as carbon nanotubes, graphene, polyhedral oligosilsesquioxane (POSS), nanofibers, nanoparticles, and high-performance engineering polymers such as polyetherimide and polyphosphazenes.

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48

MAGNUSZEWSKA, Paulina, Rafał Bogusz, and Bogdan Florczak. "STUDIES OF THE INFLUENCE OF ENERGETIC ADDITIVES ON SELECTED PROPERTIES OF HETEROGENEOUS SOLID ROCKET PROPELLANT WITH LOW CONTENT OF HCL IN COMBUSTION PRODUCTS." PROBLEMY TECHNIKI UZBROJENIA 144, no. 4 (February 27, 2018): 15–30. http://dx.doi.org/10.5604/01.3001.0011.5821.

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The paper presents influence of additives like aluminium, magnesium, AMD (aluminium-magnesium dust) and boron on selected properties of heterogeneous solid rocket propellants (HSRP) based on HTPB in which ammonium chlorate was partly replaced by sodium nitrate. The presence of sodium nitrate reduces the content of hydrogen chloride (HCl) in combustion products. Theoretical values of thermochemical and thermodynamical properties like isochoric heat of combustion (Q), specific impulse (Isp) and combustion products in motor chamber and nozzle were identified by ICT-Code program. A laboratory rocket motor (LRM) was used to examine ballistic properties for prepared samples of propellants. Their temperature of decomposition, heat of combustion and hardness were tested both with sensitivity to mechanical stimuli (impact, friction) and rheological properties at curing.

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49

Santana Jr., A., M. S. Silva, P. T. Lacava, and L. C. S. Góes. "ACOUSTIC CAVITIES DESIGN PROCEDURES." Revista de Engenharia Térmica 6, no. 2 (December 31, 2007): 27. http://dx.doi.org/10.5380/reterm.v6i2.61687.

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Combustion instability is recognized as one of the major problems frequently faced by engineers during the development of either liquid or solid propellant rocket engines. The performance of the engine can be highly affected by these high frequencies instabilities, possibly leading the rocket to an explosion. The main goal while studying combustion chambers instability, either by means of baffles or acoustic absorbers, is to achieve the stability needed using the simplest possible manner. This paper has the purpose of studying combustion chambers instabilities, as well as the design of acoustic absorbers capable of reducing their eigenfrequencies. Damping systems act on the chamber eigenfrequency, which has to be, therefore, previously known.

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50

Glebov, G. A., and S. A. Vysotskaya. "Modeling of coherent vortex structures and self-induced pressure oscillations in the combustion chamber of solid propellant." Journal of «Almaz – Antey» Air and Space Defence Corporation, no. 4 (December 30, 2016): 41–48. http://dx.doi.org/10.38013/2542-0542-2016-4-41-48.

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The study focuses on numerical analysis of the flow structure in the rocket engine chamber of solid propellant with a recessed nozzle when self-induced pressure oscillations occur. Findings of the research show that the flow structure essentially depends on the shape of the flow part of the combustion chamber and is characterized by the formation of intense toroidal vortices. We propose a method for estimating the probability of occurrence of such vortices and their effect on the intensity of pressure pulsations in the combustion chamber.

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