Solid Rocket Propulsion Research Articles

Numerical Studies of Solid Rocket Propellant PMMA-PBAN-ALF3-Nitrocellulose-Difluoramine

Abstract The aim of the new investigation is the evaluation of solid rocket Propellant. In which performing numerical analysis of solid rocket propellant and calculating the constraints involved as derived from thrust i.e. velocity involved in the solid rocket propulsion motor usage. This paper aims to show the viability of solid rocket propellant for modern and future applications. Solid rocket motors are simple in design and fabricating. The need for a propellant that produces the same specific impulse as another chemical rocket engine. So it is needed to have a solid rocket motor with an enhanced combustion characteristic at a reasonable O/F ratio. For a healthy ecology, an economically reliable, highly effective, and environmentally concerned propellant is constantly suitable. As a superior alternative among the already used propellants in the sector, the paper suggested fuel PolyMethyl MethAcrylate (PMMA), PolyButadiene AcryloNitrile (PBAN) CoPolymer, and Aluminum hydride (AlH3), and oxidizers such as Nitrocellulose and Difluoramine. The desired O/F ratio for the technical criteria was found to result in a high combustion characteristic. With the features discovered through numerical and analytical research, we have suggested a design for the SRM that includes post-combustion, which enhances the reaction and creates the innovative elements of the existing SRM.

Tetrazole-grafted-hydroxyl terminated polybutadiene: A novel energetic binder for solid rocket propulsion

The present study reports on the synthesis, characterisation, and energetic aspects of poly-nitrogen containing azole-grafted hydroxyl terminated polybutadiene (HTPB) based energetic binders (EBs). A comprehensive computational analysis at the B3LYP/6-311++G(d,p) level of theory reveals that azole-grafted HTPB variants exhibit significantly higher HoF (1061.6 to 1659.8 kJ/kg) and densities (1.26 to 1.61 g/cc) as compared to traditional non-energetic binders such as HTPB (specific heat of formation (HoF): 781.8 kJ/kg, density: 0.92 g/cc),. Their ballistic properties, including specific impulse (Isp: 266.2 to 283.8 s) and density specific impulse (ρIsp: 484.5 to 547.4 g/cc.s) are also superior as compared to HTPB (Isp: 283.8 s, ρIsp: 547.4 g/cc.s). Moreover, the synthetic methodology for grafting azole compounds onto HTPB, utilizing cost-effective, metal-free, and environmentally sustainable procedures with straightforward iodine and base mediation, underscores its potential for large-scale industrial applications. In addition to its advantages in synthesis and energetic properties, tetrazole-grafted HTPB exhibits favourable physicochemical properties: it has a low molecular weight ([Mn(VPO)] = 3023), high specific gravity (1.25 g/cc at 30 °C), moderate flowability (η = 623.6 Pa.s at 30 °C and 62.7 at 60 °C), a low glass transition temperature (−57 °C), mechanical insensitivity (low impact sensitivity and no ignition up to 400 °C), high thermal stability (onset temperature, Td,onset = 190 °C, TDSC = 246 °C), and a desirable hydroxyl value (47 mg/KOH). These properties make tetrazole-grafted HTPB an appealing choice for advancing novel propellant composites. Overall, these findings not only pave the way for advancements in HTPB-based propellants but also inspire further exploration in real-world applications, highlighting the transformative potential of azole-grafted HTPB in propellant technology.

Micro-flow properties of NEPE propellant in a rotating depressurization combustion environment of solid rocket motor

A profound understanding of microflow processes within the combustion chamber is paramount for optimizing engine performance and ensuring safe operation in the realm of solid rocket propulsion. Specifically, the scrutiny of Nitrate Ester Plasticized Polyether (NEPE) propellant under rotating depressurization combustion conditions has garnered substantial scholarly interest. The non-uniform distribution of propellant constituents and the ensuing heterogeneity in combustion rates convolve fluid flow dynamics within the chamber, ultimately engendering intricate flow patterns, turbulent phenomena, and rotational effects. Hence, conducting numerical simulation studies assumes an imperative role in comprehensively analyzing and apprehending the intricate microflow characteristics under such demanding circumstances. This paper endeavors to accomplish this overarching objective by introducing a novel kinetic mechanism in the gas phase that meticulously accounts for complicated reactions between distinct oxidizing powders and fuel binders, representing an uncharted territory in the existing literature. Additionally, this study takes into meticulous consideration the influence of radiation effects and non-planar moving surfaces, as well as other pivotal components. Upon establishing the combustion model, exponential depressurization is induced within the gas phase, while the computational domain is deflected to accurately simulate the rotational effects. A series of methodical investigations is subsequently undertaken to validate the proposed numerical framework, yielding commendable predictive accuracy. Ultimately, this research meticulously examines the minute intricacies of microflow characteristics under such complex transient combustion conditions.

EFFECTS OF FLUORINATED AND NON-FLUORINATED ADDITIVES ON BURNING RATES OF BORON-TEFLON BLENDS FOR SOLID ROCKET PROPULSION

Boron is considered a promising metal fuel for energetic propellants due to its high energy density, but in practice, boron experiences slow combustion kinetics. Fluorinated additives can improve boron combustion. The objective of this study was to determine how different fluorinated additives would affect the burn rate of a boron-based fuel in a high-pressure inert environment. Boron powder was sintered with polytetrafluoroethylene (PTFE) powder to create test pellets with variable fluorine:boron mass ratios. Three different purities of boron powders were tested, along with three different fluoropolymer additives and 14 other additives, which varied in molecular morphology and fluorination. The base boron-PTFE test pellets with no additives have a maximum burning rate at fluorine:boron mass ratio of ~ 3.5, which remains the same even with additives at 20 wt%. A commercial fluoropolymer, Aquivion, showed significant burn rate enhancement at 20 wt% but not at 5 wt%. At this lower loading level, materials that tend to enhance thermal conductivity (graphene, nanotubes) show some burn rate enhancement, even without fluorination. However, very significant burn rate enhancement is shown by a fluorinated carbon nanotube, likely due to the reduced bond dissociation energy of fluorine when attached to a nanotube. The data suggest this method of fluorination can have an important effect on enhancing boron combustion, and even small additives with readily available fluorine could significantly improve boron combustion rates for solid propellant applications.

ANALYSIS OF FOREIGN EXPERIENCE OF DUAL USE SOLID ROCKET MOTORS

In countries, such as the United States, China, India, France, Israel and many others, solid-propellant rocket engine technologies are used not only in military missiles, but also in launch vehicles of the rocket and space industry. Such an approach of dual application of technologies makes it possible to introduce and develop new innovative solutions to increase the degree of unification of related technologies. One of the most important factors influencing the efficiency of using solid propellant rocket engines in space programs is the cost of solid rocket propellants and propulsion system components. For this reason, a high degree of unification of solid fuel products for various purposes is aimed at ensuring a high utilization of the industry’s production capacities, which can significantly reduce prices for materials, fuel and the design of propulsion systems as a whole. The analysis confirmed that the creation of launch vehicles based on rocket engines of solid fuel for military missiles naturally leads to a reduction in the time and costs for the development of such products, a reduction in technical risks and an increase in the reliability of the technologies used.

FOX-7 derived nitramines: Novel propellants and oxidizers for solid rocket propulsion

A variety of amine derivatives was found to react with FOX-7 to give triazine derivatives, which upon nitration gave rearranged nitramines in good to excellent yields. The structures of these triazines and nitramines were determined by various spectroscopic techniques, and elemental analyses. Further the structure of selected compounds namely 4-(dinitromethylene)-1-methyl-1,3,5-triazinane (8a), 4-(((methyl(nitro)amino)methyl)imino)-3,5-dinitro-5,6-dihydro-4H-1,2λ4,5-oxadiazin-2-olate (9a) and 3,5-dinitro-4-(((nitro(2-(nitrooxy)ethyl)amino)-methyl)imino)-5,6-dihydro-4H-1,2λ4,5-oxadiazin-2-olate (9b) were unambiguously confirmed by single crystal X-ray analysis. These compounds have good thermal stabilities with their decomposition temperatures in the range of 119–188 °C and experimental densities between 1.38 g cm−3 to 1.75 g cm−3. The enthalpies of formation of the new compounds were calculated (Gaussian-03) and combined with their experimental densities to evaluate the propulsive and detonation properties. Compounds 9a and b are found to have excellent energetic properties (9aP = 31.48 GPa, D = 8542 m/s; 9b, P = 31.83 GPa, D = 8518 m/s) that exceed those of TNT, HNS and are comparable to FOX-7 and RDX. The potential of these compounds as solid propellants/oxidizers was explored in detail; compounds (9a, Isp,vac 290.74 s; 9b, Isp,vac 293.00 s; with AP, Isp,vac 291.66 s (9a) and Isp,vac 292.57 s (9b)) were found to be suitable oxidizers for solid rocket propulsion.

Design and Synthesis of Nitro Cage Heterocycles as Energetic Materials Derived from Pentacycloundecane (PCUD) Systems

AbstractA variety of pentacycloundecane (PCUD) based cage compounds containing nitro groups were synthesized via a simple synthetic method starting with cage diones. Eight PCUDs were used for nitro functionalization. These cage diones were prepared starting with easily accessible starting materials such as 2,3‐dimethylhydroquinone, 1,4‐dihydroxybenzene, and cyclopentadiene. Here, we have used Diels‐Alder (DA) reaction, [2+2] photocycloaddition, and Henry reaction as key steps. Transannular cyclization plays an important role to generate a nitro substituted cage heterocycles via base promoted reaction using nitromethane as a reagent. Some of these oxa‐cage structures have been further supported by single‐crystal X‐ray diffraction studies. Energetic performance of the title compounds was evaluated using DFT based calculations. These bis(nitromethyl)octahydro‐1H‐1,3,4,6(epiethane‐[1,1,2,2]tetrayl)‐pentaleno[1,6‐bc]furan analogues (11 b–18 b) owned the densities in the range of 1.37–1.60 g/cm3. All the studied compounds possess high positive enthalpy of formation (33.68–230.48 kJ/mol, except compound with dimethyl groups (16 b, −11.38 kJ/mol) and very good thermal (Tm=99–147 °C) and kinetic stabilities. Ammonium perchlorate based various propulsion formulations suggests that these compounds could be excellent candidates in the solid rocket propulsion.

Application Energetic Materials for Solid Composite Propellant to Support Defense Rocket Development

In its application in space technology, solid composite propellants are often used as fuel in rockets for military purposes. Increasing the energy of the propellant is carried out by observing two stages, the use of energetic materials and improvements to the process technology. The current development of propellant technology makes it possible to use new energetic materials, simple formulations, high energy, and smokeless. The purpose of this research is to find out developments related to the use of highly energetic materials as raw materials for composite propellants for defense rockets at the Rocket Technology Research Center, ORPA-BRIN. This study uses qualitative analysis methods with research designs in the literature studies and simulation results. In the context of mastering rocket propellant technology in Indonesia, the application of highly energetic materials is expected to be able to solve the problem of rocket propulsion performance. Currently, the Rocket Technology Research Center, ORPA-BRIN is developing a smokeless propellant composite with a composition based on the energetic materials AP/HTPB/Al and an oxidizing agent RDX. From the results of the combustion simulation software ProPEP and RPA, it shows that the composition of the resulting combustion gaseous (Al2O3 and HCl) shows a decrease when using RDX energetic material-based propellant. It's known that RDX can significantly reduce smoke in propellant combustion products. The application of the new highly energetic materials compound is expected to significantly solve the problem of solid rocket propulsion performance.

Energetic performance of trinitromethyl nitrotriazole (TNMNT) and its energetic salts.

Little is known about trinitromethyl nitrotriazole (TNMNT) since the crystal structure, density, energetic performance, and thermal properties have not been determined. A detailed characterization of TNMNT and its hydrazinium and potassium salts and their potential as solid propellants and oxidizers has been established. TNMNT exhibits a high density (1.96 g cm-3) and positive enthalpy of formation (ΔHf = +84.79 kJ mol-1). TNMNT and its hydrazinium and potassium salts illustrate excellent detonation properties (P = 34.24 to 36.22 GPa, D = 8899 to 9031 ms-1). TNMNT and its hydrazinium salt exhibit outstanding propulsive properties (Isp = 247.28 to 271.19 s), and these are superior to AP (Isp = 156.63 s) and ADN (Isp = 202.14 s). The results suggest opening the door to utilizing TNMNT and its energetic salts in solid rocket propulsion.

Trinitromethyl-triazolone (TNMTO): a highly dense oxidizer.

A scalable synthesis of 5-(trinitromethyl)-2,4-dihydro-3H-1,2,4-triazol-3-one (TNMTO) is possible from commercially available 2-methylpyrimidine-4,6-diol. It exhibits high density (1.90 g cm-3) with comparably low thermal stability (Td = 80 °C) and positive oxygen balance (OBco = 20.51%, OBCO2 = 0.0%). TNMTO has an attractive combination of detonation properties (P = 35.01 GPa, D = 8997 ms-1) and propulsive properties (Isp(neat) = 251.85 s, ρIsp(neat) = 478.52 gs cm-3, ). These are superior to ammonium dinitroamide (ADN), 2,2,2-tetranitroacetimidic acid (TNAA) and ammonium perchlorate (AP), making it a potential green oxidizer in solid rocket propulsion.

Design Optimization of Propellant Grain and Nozzle Contour to Improve Performance of Solid Rocket Propulsion

A rocket is a spacecraft, guided missile, or flying vehicle that boosted by a chemical reaction resulting from the combustion of propellant in the rocket motor. One of the essential parameters in the development of rocket motors is design optimization to improve the propulsion performance of the rocket. Increasing the propulsion performance of the rocket will increase the flight performance of the rocket, in terms of its maximum range or the altitude of the rocket trajectory. This study examined the determination of the design parameter values of a rocket motor by looking at it as an optimization problem with constraints. The problem studied was limited to the case of the second-stage rocket motor. A genetic algorithm was used to solve the resulting optimization problem of propellant grain configuration cases and a characteristic method for designing the bell nozzle. The results obtained indicated an increase in total impulse by 10% compared to the results before optimization.

Thermal interactions between hybrid HMX/ANPyO cocrystals and commonly used propellant ingredients

Nitramine-based composite propellants (N-CPs) are receiving significant attention in the field of solid rocket propulsion. Cyclotetramethylene tetranitramine (HMX)/2,6-diamino-3,5-dinitropyridine-1-oxide (ANPyO) cocrystals are a novel insensitive energetic material that may replace the high-energy component HMX in insensitive formulations. This study used the DSC-TG coupling technique and the in-situ FTIR technique to investigate the effects of propellant ingredients (AP, n-Al, and nano-Fe2O3) on the thermal decomposition of hybrid HMX/ANPyO. The results show that there are strong interactions among the components of these mixtures. In particular, the peak thermal decomposition temperatures of HMX/ANPyO/AP decreased by about 38.1–42.6 ​°C. In addition, the activation energy (Ea) of hybrid HMX/ANPyO cocrystals was reduced under the effects of Al and nano-Fe2O3. The hybrid HMX/ANPyO cocrystals have different flame propagation speeds, whereas the two typical binary mixtures of HMX/ANPyO/nano-Fe2O3 each have the lowest flame propagation speeds of 0.9 ​mm·s−1 and 1.0 ​mm·s−1 at ambient conditions. The flame propagation speed of HMX/ANPyO/n-Al was barely dependent on pressure, but those of HMX/ANPyO/AP and HMX/ANPyO/nano-Fe2O3 decreased with an increase in pressure.

Observation of aluminum interaction with the binder melt layer using high-speed synchrotron-based phase contrast imaging

Particle interactions with the binder melt layer are a major factor in the combustion efficiency and stability of the propellant in a solid rocket motor. Metal particles tend to agglomerate on the surface of burning solid propellants, inhibiting the combustion process. Therefore, reduction of molten aluminum agglomeration in a solid rocket motor is vital to the improvement of solid rocket propulsion system performance. In other work, quenched samples have been used to study the impact of how metal particles alter the flow properties of the molten propellant surface. In-situ optical measurements have also been attempted for these particle-condensed phase interactions, but with little success due to the opacity and strong gradients within the flame. This study expanded on previous work using synchrotron–based x-rays to directly image these aluminum agglomerates as they interacted with the binder. X-ray absorption and phase contrast in the images allowed for the direct measurement of the particle size in combusting propellants in-situ at typical rocket pressures. Particle size distributions were collected in an optically accessible combustion vessel over a pressure range of 1.4- to 6.9-MPa (200- to 1000-psig). This investigation measured the aggregate sizes ranging from 300-to 650-µm at a 6.9 MPa chamber pressure. Interestingly, the lowest binder melting temperature produced the longest aluminum chaining at 1.1 mm and the highest binder melting temperature produced the second longest chain at 524 µm long. The study observed aluminum interactions with the binder melt layer on the propellant surface that may be a contributing factor in a plateau propellant. Additionally, data pointed to possible relationships between binder viscosity, burning rate, aluminum clusters, and chain formation in a solid propellant. The study concluded with collecting data on the molten aluminum volumetric changes in a propellant environment, thus defining a starting particle diameter for particle regression rates. Collecting the regression rates also provided information on the evaporation rate of aluminum, which burned in the propellant flame. All of these data are critical for understanding the combustion of aluminum-based solid rocket propellants and outlines knowledge gaps that need additional data.

Hydrogen peroxide – A promising oxidizer for rocket propulsion and its application in solid rocket propellants

This paper discusses the potential use of hydrogen peroxide as oxidizer for solid rocket propulsion. While hydrogen peroxide is a liquid in normal conditions, it may be used in solid rocket motor grains. Use of hydrogen peroxide of HTP class (High Test Peroxide) is proposed. It has been known for many decades but its utilization has been historically limited to liquid propellants. Until recently is application even in liquid state in rocket propulsion for space and defence has been limited due to storability and safety issues. With the availability of new grades of HTP with higher purity and concentrations, enhanced performance, safety and storability are possible. This results in HTP being considered for a wide range of rocket propulsion systems: as oxidizer in bipropellant and hybrid propulsion systems, as well a monopropellant. To ensure proper analysis of the potential of using HTP in next-generation solid rocket propellants, this paper reviews existing rocket propulsion applications of HTP. Modern use requires high performance and common current composite propellant compositions utilize ammonium perchlorate (AP) as oxidizer, what has several disadvantages, which are discussed. Alternative oxidizing compounds include ammonium dinitiramide (ADN), ammonium nitrate (AN), hydrazinium nitroformate (HNF), hexanitrohexaazaisowurtzitane (HNIW) and several other secondary explosives. Key properties of solid propellant oxidizers are listed and discussed. The need for further alternatives, despite numerous recent advances in solid rocket oxidizer technology is justified. The global push forward high performance green propulsion is one of the main motivations behind considering HTP for solid rocket propulsion. Theoretical performance of solid rocket motors using solid grains containing a high mass fraction of HTP is presented. This includes performance considering several fuels and additives and different oxidizer loadings. Up to date concepts of using hydrogen peroxide in solid propellants are reviewed. Solid cryogenic propellants using hydrogen peroxide are mentioned, but focus is given to solid propellants which could ensure flexible operations, thus use in state-of-the-art solid rocket motors. Challenges of HTP application as an oxidizer for solid propulsion are listed. This includes a discussion concerning its reactivity, thus limited compatibility with organic materials. Recommendations for further experimental work are proposed. Potential technology applications are listed with explanations why these particular niches may benefit from utilization of the new solid propellant technology.

Optical Diagnostics for Solid Rocket Plumes Characterization: A Review

In recent decades, solid fuel combustion propulsion of spacecraft has become one of the most popular choices for rocket propulsion systems. The reasons for this success are a wide range of applications, lower production costs, simplicity, and safety. The rocket’s plumes leave the nozzle at high temperatures; hence, the knowledge of produced infrared (IR) emissions is a crucial aspect during the design and tests of the rocket motors. Furthermore, rocket plume composition is given by N2, H2, H2O, CO and CO2, while solid rocket motors (SRM) additionally inject some solid particles, given by metal fuel additives in the propellant grain, i.e., aluminum oxide (Al2O3) particles. The main issue is the detection of the particles remaining in the atmosphere due to the exhaust gas of the solid rocket propulsion system that could have effects on ozone depletion. The experimental characterization of SRM plumes in the presence of alumina particles can be conducted using different optical techniques. The present study aims to review the most promising ones with a description of the optics system and their potential applications for SRM plume measurements. The most common measurement techniques are infrared spectroscopy imaging, IR imaging. UV–VIS measurements, shadowgraph, and Schlieren optical methods. The choice of these techniques among many others is due to the ability to study the plume without influencing the physical conditions existing in and around the study object. This paper presents technical results concerning the study of rocket engines plumes with the above-mentioned methods and reveals the feasibility of the measurement techniques applied.

A TREATISE ON NANOALUMINUM FOR SOLID ROCKET PROPULSION

A TREATISE ON NANOALUMINUM FOR SOLID ROCKET PROPULSION

Correction to: Ferrite Nanoparticles: Synthesis, Characterization, and Catalytic Activity Evaluation for Solid Rocket Propulsion Systems

The original version of this article unfortunately contained a mistake. The authors reused an image taken from a book without referencing it. The missing reference is given below: Zarko, V. E. and A. A. Gromov, Eds. (2016). Energetic Nanomaterials Synthesis, Characterization, and Application. Amsterdam, Elsevier. p 325.

Dual-sheet interferometric particle imaging for opaque particle size and 2D location measurement

Dual-beam interferometric particle imaging (DIPI) has been proposed to measure opaque spherical particle size in the single-point region of interest. This work extends DIPI to measure the size and two-dimensional position of opaque spherical particle in a plane with dual-sheet illumination. A general DIPI model based on geometric optics is established to formulate the interference signal formation of DIPI by modeling the phase difference of two reflected light at far-field. Proof-of-concept experiments, incorporating digital inline holography (DIH), are performed for validation. An interferogram processing algorithm which adopts the sum-of-squared-difference image registration method and Fourier spectrum analysis has been proposed to locate the centroid of each out-of-focus pattern and extract the spatial frequency of fringes, and subsequently particle position and size. The average relative deviation values of DIPI in size measuring and two-dimensional locating are 2.4% and 1.9%, respectively, compared to those values by DIH, which demonstrates the practicability of DIPI in both respects. The extension of DIPI from point-probe to planar measurement promotes its applications to real scenarios, e.g., metal droplet in solid rocket propulsion.

Characterization of SRM plumes with alumina particulate in subscale testing

The current paper provides an outline and first results of the ESA-EMAP project. This project pursues activities regarding the experimental modeling of alumina particulates in solid boosters (EMAP). The issue regards the particles residing in the atmosphere after the passage of a launch vehicle with solid rocket propulsion, which might contribute to local and overall ozone depletion. The question is to what extent since the particle size distribution left behind is essentially unclear. For this reason, the ESA-EMAP investigations focus on the characterization of the solid exhaust plume properties for well-defined combustion chamber conditions. Thus, details of the rocket motor assembly, of the developed solid propellant grains, and of first measurement results are provided. The paper presents technical findings concerning the rocket motors and reveals aspects to the feasibility of the applied measurement techniques.

Design of a Solid Rocket Propulsion System

Design of a Solid Rocket Propulsion System