Solid Rocket Propellants Research Articles

Numerical simulation study of aluminum particle evaporation combustion process based on SPH method

A large variety of metal fuels is usually added to the modern solid rocket propellants to improve the propellant energy and motor specific impulse and to suppress high frequency unstable combustion. Among them, aluminum is the most common additive. The combustion process of aluminum can significantly affect the combustion characteristics of a solid rocket motor. In this paper, the SPH (Smoothed Particle Hydrodynamics) method is used to simulate the combustion process of aluminum particles. First, the SPH discrete equations with an evaporation combustion model are derived. On this basis, the evaporation combustion process of aluminum particles is numerically simulated. The results show that when aluminum particles are heated and evaporated in a static flow field, an alumina shell will be formed on the surface, and further thermal expansion will cause the alumina shell to break. The molten aluminum will spray out, and an aluminum cap will be formed on the surface. The “microburst process” will be similar to the gel droplet. In the convective environment, the flame structure of aluminum particles will be obviously peach shaped, which will wrap aluminum particles at the bottom. The simulation results are consistent with the experimental results.

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.

First applications of ultrasound technology in solid rocket propellant combustion promotion

The regulation of propellant combustion using ultrasonic waves is proposed. Under ultrasonic frequencies of 25–40 kHz, we systematically examined the combustion characteristics of Al particles, ammonium perchlorate (AP)/hydroxyl-terminated polybutadiene (HTPB) propellants, and Al-containing Al/AP/HTPB propellants. Ultrasonic treatment increased the ignition delay time of the Al particles by 48.3 %. However, it increased the burning rate of the AP/HTPB propellants by up to 26.1 % and decreased their ignition delay by 39.3 %. As the ultrasonic frequency increased, the burning rate of the solid Al/AP/HTPB propellants increased by 22.5 %, and the degree of Al agglomeration decreased, resulting in a 24 % decrease in the size of the condensed-phase combustion products. The action mechanism was analyzed in terms of the effects of ultrasonic waves on Al droplets and combustion flames. The introduction of ultrasonic waves split the Al droplets near the burning surface into smaller particles, which affected combustion efficiency. Bringing the diffusion flame closer to the burning surface affected the burning rate. These findings demonstrate that the combustion and agglomeration characteristics of solid propellants can be modified using ultrasonic techniques, providing a new method for controlling the thrust of solid rocket motors.

Insight into HTPB pyrolysis mechanism under high-temperature: A reactive molecular dynamics study

Hydroxyl-terminated polybutadiene (HTPB) is widely utilized in solid and hybrid rocket propellants due to its mechanical properties and combustion performance. Insight into its pyrolysis process is key to enhancing combustion efficiency in rocket engines. This study employs ReaxFF molecular dynamics (MD) simulations to explore the pyrolysis mechanism of HTPB under extreme conditions, with temperatures ranging from 1000 K to 2000 K. The results identify the primary degradation products and elucidate their formation mechanisms. The simulation reveals that C-C bond cleavage at polymerization sites is the initial step, followed by the formation of linear oligomers and butadiene. Subsequent reactions, including hydrogenation and dehydrogenation, lead to the generation of smaller molecular species. The kinetic analysis confirms that HTPB pyrolysis follows first-order reaction kinetics, with an activation energy of 8.12 kcal/mol. The findings are compared with existing experimental data, highlighting the influence of thermal environments on pyrolysis mechanisms and product distributions.

Combustion and mechanical properties enhancement strategy based on stearic acid surface activated boron powders

Boric acid and other impurities on the surface of boron (B) particles can interact with hydroxyl-terminated polybutadiene (HTPB), weakening the mechanical properties and energy release efficiency of boron-based solid rocket propellants. SA@B composite particles were created by coating stearic acid (SA) on the surface of B particles through solvent evaporation-induced self-assembly. The study investigated the impact of SA coating on the combustion performance of B particles and the mechanical properties of HTPB matrix composites. The results showed that the SA coating enhanced the oxidation efficiency of B particles in air. The combustion heat of SA@B composite particles is 30.29 MJ/g, about 50% higher than that of B particles. During the combustion of SA@B composite particles, fewer molten solid particles surround the flame, which enhances the stability of the combustion process of the B particles. Furthermore, the SA coating effectively enhanced the dispersion of B particles in HTPB. At a stretching speed of 100 mm/min, the tensile strength of the SA@B/HTPB composite materials is higher than that of the B/HTPB composite materials. Moreover, when the mass loading of the SA@B composite particles reaches 50 wt%, the tensile strength of SA@B/HTPB composite materials is 2.46 MPa. Activating the surface of boron particles with SA can significantly improve their compatibility with HTPB, which is crucial for the stable storage of boron-based solid rocket propellants.

Numerical and experimental study on the curing process of NEPE solid rocket propellant considering multi-physical field effects

A multi-physical field numerical model considering the thermo-chemical and mechanical effects was used to simulate the curing process of a nitrate ester plasticized polyether (NEPE) solid propellant. The curing process of solid rocket propellants was detected through physical experiments such as differential scanning calorimetry (DSC) and heat transfer. The variation laws and distribution patterns of the temperature field and curing degree field during the propellant curing process were numerically and experimentally analyzed, and the residual and contact stresses caused by the effects of temperature, chemical reaction, and pressure were studied. The results showed that the internal temperature field of the NEPE propellant during the curing process was closely related to the curing time. At about 3 d of curing time, the propellant curing rate was the fastest, and the temperature reached a maximum value of 58 °C. During curing, the NEPE solid propellant produces an unevenly distributed temperature field and curing degree field distribution, which aggravates the residual stress. The molding pressure can reduce the residual stress; however, when the pressure exceeds approximately 3 MPa, it causes extrusion stress on the interface layer between propellant-to-case and propellant-to-core mold. Based on the numerical calculation results, an optimal pressure calculation formula for multiple parameters that can quickly estimate the optimal pressure is derived.

Mechanical properties assessment of composite solid propellant based on different aziridine bonding agents

Abstract The employment of the dynamic mechanical analyzer (DMA) represents a modern methodology for the monitoring of the thermo-mechanical characteristics and also the thermomechanical loads of polyurethane-based composite solid rocket propellants during combustion. A meticulous preparation of composite propellant samples based on different bonding agents was carried out to examine the influence of various bonding agents on mechanical properties. To simulate the effects of natural aging, an accelerated aging program was implemented and subjected to the propellant samples for 5-day exposure at 90°C, equivalent to 5 years of natural aging. The dynamic mechanical modern analysis technique was used to assess the viscoelastic behavior and degree of viscoelasticity of both fresh and aged propellant samples. It is concluded that M4 based on MT-X bonding agent is the most valuable formulation of CSRP with tan δ of 0.183591, also with mechanical results; implying that M4 formulation based on MT-X bonding agent has the highest degree of viscoelasticity.

EVALUATION OF SOLID ROCKET PROPELLANTS FOR LOW EARTH ORBIT

In contemporary aerospace technology, solid rocket propellants are commonly employed in the thriller and boost of space cargoes to LEO. These are reliable, uncomplicated, and high in thrust; and thus, can be used in both military as well as civil space ventures. This paper provides a comprehensive evaluation of five prominent solid rocket propellants: Five categories namely Ammonium Perchlorate Composite Propellant (APCP), Double-Base Propellant, Composite Modified Double Base (CMDB) Propellant, Hydroxyl-terminated polybutadiene (HTPB)-Based Propellant, and Ammonium Nitrate Composite Propellant (ANCP) have been identified. This also falls under the evaluation and includes aspects such as historical bases to see how the formulation of the propellant has changed over time regarding the development of propellant technology. In the case of each propellant, the chemical properties of the mix and reactions that its fundamental are analyzed. It also shares details regarding the changes and advancements made in the little while to enhance capability, safety, and environmental compatibility of these fuels, such as synthesizing fuels for liquid bipropellants. Some of the modern applications of technologies within the active spectrum of aerospace operations are explained in more detail to provide clients with the most versatile and essential information. Among the essentials of this work, the complex calculation of thrusts developed by the specific propellant with corresponding graphics has significant importance. These estimations afford chances of comparing the efficiency of the propellant in terms of producing thrusts as well as achieving the indicated optimal performance. Thus, the main objective of this study is to identify which of the solid rocket propellants to use in LEO missions and if any, what modifications could be made to improve them. Thus, the purpose of this paper is to contribute to the development of the constantly evolving branch of rocket propulsion by discussing the benefits and shortcomings of each propellant. These findings and recommendations could be useful in perfecting the recipes for the propellants and improving the efficiency of the subsequent missions to planets. Keywords: Aerospace, Solid rocket propellants, Thrust, Low earth orbit.

Comparative Analysis of Solid Propellant Rocket Fuel Efficiency: Gunpowder vs. Sorbitol and Potassium Nitrate

Comparative Analysis of Solid Propellant Rocket Fuel Efficiency: Gunpowder vs. Sorbitol and Potassium Nitrate

Impact of Solid Rocket Propellant Grain Manufacturing Limitations on Launch Vehicle Capability

It is examined if any limitations in existing solid rocket propellant grain manufacturing methods adversely affected the payload capability of recent space launch vehicles. It is seen if the transition from heavy, segmented metal rocket motor casings to lightweight monolith composite casings is possible without loss of ability to design and realize high-performance grain configurations using simple and safe methods. Considering payload fraction as the comparative performance metric, recently flown solid rocket-propelled, small-lift launch vehicles were surveyed and ranked. Solid rocket boosters of underperforming launch vehicles were investigated for manufacturing factors influencing payload fraction by comparing them to boosters of better-performing launch vehicles in their weight class. Relationships between payload fraction and the solid boosters’ mass fractions, casing construction, shape of thrust profile, propellant grain configuration and method employed to manufacture the grain were analysed. It is shown that those launch vehicles that did not possess or use the technology necessary to manufacture high-performance grain configurations like undercut finocyl in monolith composite casings ended up having boosters delivering poor thrust profiles with high inert mass ultimately leading to low payload fractions.

Study of Aging Characteristics for Metalized HTPB Based Composite Solid Propellants Stored in Ambient Conditions

The aging of any propellant is defined as the change in the physical, chemical, and performance parameters of solid rocket propellants. The propellant’s service life and aging properties are important parameters of the study, especially for missiles and other defense applications. Hydroxyl-terminated polybutadiene (HTPB) based composite solid propellants with ammonium perchlorate (AP) are the most prominently used propellants in the operations of solid rocket motors in the defense and space sectors. Thus, studying this composite solid propellant is of essential when determining ambient service life. Performance parameters studied in this research are burn rate under high-pressure conditions in Crawford bomb setup, Thermogravimetric Analysis, and Fourier Transform Infrared Spectroscopy (FTIR). SEM and X-ray diffraction (XRD) analysis of the aged sample were also conducted to ascertain the chemical composition and morphological changes in the samples. Naturally aged propellant strands manufactured in different years have been compared with freshly prepared ones to establish a trend for deriving conclusions. The results from different analysis techniques, FTIR, XRD, and FESEM, depicted that oxidation of metals happens while aging of propellant due to atmospheric moisture, and the metal oxides prominently affect the propellant chemical composition and decomposition process of the propellant samples. The ballistic properties of the aluminium added samples showed an increment in burn rate. In contrast, the bimetal addition of aluminium and magnesium combined as an additive decreased the ballistic burn rate.

Design of 2-DOF Thrust vector control system for Rockets and Missiles

This paper outlines the development and implementation of a Thrust Vector Control (TVC) system tailored specifically for solid propellant rocket engines, aimed at enhancing the portability of hybrid and liquid propellant rocket systems. The proposed system ensures trajectory stability and rapid response to flight emergencies by effectively addressing external perturbations like wind. Key components are Gyroscopic (GYRO) sensors, good micro-controller, and Servo motors, collectively referred to as Thrust vectoring components, which dynamically adjust thrust direction opposite to the rocket's trajectory to maintain stability. Drawing from an extensive literature review analyzing existing TVC system designs, with a focus on thrust characteristics and engine combustion timings, the design integrates considerations of geometric properties, kinematics, forces, energy requirements, safety, cost, and control methods, aligning with both mechanical and avionic design principles. The anticipated outcome is an advancement in rocket propulsion technology, characterized by improved angular speed control, trajectory linearity, and rapid response capabilities.

Моделювання інжекції газу в надзвукову частину сопла при газодинамічному регу-люванні вектора тяги

Solid-propellant rocket engines are simple in design, highly reliable, and able to store the propellant for a long time without its degradation. Their main feature is that the propellant is a mixture of a solid fuel and a solid oxidizer, thus ensuring a uniform combustion and a stable discharge of the combustion products. However, the combustion rate cannot be controlled, and the combustion cannot be stopped or restarted. This calls for efficient methods of thrust vector control. Gas-dynamic methods, such as a gas injection into the supersonic nozzle area, offer a required flight path control without complex high-power mechanical systems. The importance of this study lies in improving the accuracy and efficiency of rocket flight control, which is critical for today’s space and defense tasks. The numerical simulation of gas-dynamic control systems, in particular by an asymmetric gas injection, allows one to obtain detailed data on the flow behavior and optimize the design and operating conditions of the system. This study is concerned with a full-scale solid-propellant rocket engine with a gas-dynamic thrust vector control system based on the use of asymmetric forces that occur on the nozzle wall when the supersonic flow interacts with the injected transverse jets. To simulate the process in the Ansys Fluent software package, a geometric model of a nozzle with an asymmetric injection of the chamber gas into the supersonic area was developed. The injection flow rate was controlled by moving the valve flap. The simulation was carried out taking into account the temperature dependence of the main thermophysical gas parameters with consideration for dissociation processes by way of data approximation. The approximation was performed using piecewise polynomial functions. Nozzle flow patterns were obtained. The calculated results were compared with experimental test data and shown to be in satisfactory agreement with the lateral force measured during the fire bench tests of the prototype. From a practical point of view, the results obtained may be directly used to improve existing thrust vector control systems and develop new ones. This will improve rocket navigation accuracy, flight stability, and maneuverability, which is critical for complex space and defense tasks.

Dynamic mechanical response and failure behavior of solid propellant under shock wave impact

Solid rocket motor propellants are typically subjected to shock wave loads during ignition. To analyze the dynamic mechanical response and failure behavior of hydroxyl-terminated polybutadiene (HTPB) propellant under varying shock intensities, this study innovatively employed a shock tube apparatus in conjunction with schlieren imaging and 3D-DIC techniques. Shock loading experiments were conducted on propellant samples of three different thicknesses under pressures ranging from 0.3 to 0.7 MPa, capturing the dynamic deformation processes. Results revealed that deformation initiated at the clamped region, extending towards the center and forming a parabolic shape. The maximum deformation of the samples shows an approximately linear relationship with the launch pressure and decreases with increasing sample thickness. The 5 mm sample failed under a 0.6 MPa shock pressure, with electron microscopy images identifying matrix damage, particle debonding, and particle fracture as the principal failure mechanisms. Finite element simulations based on the generalized Maxwell linear viscoelastic constitutive model were conducted, demonstrating good agreement with experimental data. The critical stress for fracture was determined to be 4.12 MPa, and the ultimate shock wave pressures for 3, 6, 9, and 12 mm thicknesses were approximately 0.3 MPa, 0.65 MPa, 1.1 MPa, and 1.4 MPa, respectively. These findings provide valuable insights for assessing the structural integrity of solid rocket motors under ignition shock conditions.

Analysis of the mechanical response of solid rocket engine propellant under acceleration shock

The increased-range solid rocket engine of new ammunition often needs to withstand extremely high acceleration overload, leading to damage to the propellant’s structural integrity. To address this issue, this paper constructs a nonlinear viscoelastic constitutive model for the CMDB propellant by using low- and high-strain rate mechanical property tests. Combined with the secondary development of explicit dynamics and numerical simulation technology, the mechanical response of a certain type of solid rocket propellant under different acceleration impacts is analyzed. The results show that under acceleration impact loads, the propellant will compress, rebound, and recover over time, continuously cycling. Its axial displacement, maximum equivalent stress, and maximum equivalent strain exhibit irregular sinusoidal wave-like periodic cycles. Looking at the time when the peaks appear, the time when the maximum equivalent stress appears always lags behind the time when the maximum axial displacement peak appears. Due to the viscous effect of the viscoelastic material of the propellant, the time when the equivalent strain peak appears will lag behind the equivalent stress. Because of the material’s damping effect, both the peak values of the maximum equivalent stress and equivalent strain decrease over time. Under continuous high acceleration impact loads, this viscous damping phenomenon continuously diminishes, and the peak value of the propellant’s axial displacement gradually increases.

Transient plasma enhanced combustion of solid rocket propellants

We demonstrate the use of nanosecond pulse transient plasma (NPTP) to improve the control (and acceleration) of the combustion of solid rocket propellants. Here, we fabricate end-burning propellant samples (i.e., grains) with a co-axial center wire electrode using hydroxyl terminated polybutadiene (HTPB), isodecyl pelargonate (IDP), modified diphenyl diisocyanate (MDI), and ammonium perchlorate (AP) as the fuel, plasticizer, curative, and oxidizer, respectively. High voltage (20 kV) nanosecond pulses (20 nsec) produce a streamer discharge that provides electronic throttling of the solid rocket propellant. These studies are carried out over a wide range of oxidizer mass fractions, including those considered insensitive munitions (IM). In addition, real time imaging is performed characterizing the plasma-formation, evolution of the ignition process, and plasma enhanced flame-fuel coupling. We believe the plasma-based mechanisms of enhancement are 3-fold: 1.) The plasma provides highly energetic electrons that drive new chemical reaction pathways via highly reactive atomic species such as H, O, and Cl, 2.) The plasma sputters chunks of the solid fuel material up into the flame where it is combusted, producing an agitated flame profile. 3.) The plasma provides increased turbulence and multi-scale mixing due to hydrodynamic effects (i.e., ionic winds), which further improves the combustion process. Having electronic control of the burn rate introduces the ability to “throttle” solid rocket motors and introduce new flight profile options beyond a pre-selected profile such as the typical boost-sustain profile. While we are unable to quantify the burn rate or thrust from these relatively simple observations, we observe clear evidence of the effect of the plasma on the combustion of these solid rocket fuels even at high oxidizer content.

Application of co-crystallization method for the production of ammonium perchlorate/ammonium nitrate oxidizer for solid rocket propellants

Efforts in the quest for greener alternatives to ammonium perchlorate (AP) as a solid rocket propellant oxidizer have intensified due to the associated health and environmental concerns. AP exceptional performance in propulsion is accompagnied by its environmental concerns, necessitating the exploration of greener alternatives. Ammonium nitrate (AN) has emerged as a promising substitute, offering cleaner combustion products and cost-effectiveness. AN-based propellants face challenges such as hygroscopicity and structural instability. This study aims to reconcile the latter challenges by employing a co-crystallization process that embodies a synergistic association of AP and AN, culminating in a novel molecule that harmonizes their distinct advantages and mitigates their individual shortcomings. Optimizing the solvent/antisolvent co-crystallization process through a variation of the antisolvent temperature Ta and antisolvent-to-solvent ratio Ra/s allows to assess their impact on the co-crystallization process of AP and AN and to attain the desired and optimal co-crystal formation conditions. Various characterization techniques were applied to gain deep insights into the co-crystal formation mechanism, its thermal behavior, morphological and structural features together with its thermal decomposition kinetics. Maintaining a low Ta and a high Ra/s (Ta = 15 °C, Ra/s = 11) yielded the design of AP/AN that demonstrates typical indications of the cocrystal formation. The AP/AN cocrystal formation elevated the decomposition temperature by about 18 °C compared to the physical mixture. Furthermore, the combination of AP and AN demonstrates a non-negligible decrease in the HTD temperature compared to that of pure AP by about 110 °C. The kinetic study demonstrated a 50 kJ/mol increase in the activation energy of the AP/AN cocrystal compared to that of the physical mixture. The present research addresses the critical gap in the literature by investigating the co-crystallization of AP and AN that not only improves the safety and environmental impact of solid propellants but also enhances their overall performance and stability.

Titanium a novel catalyst and high energy dense material: an opportunity for composite propellants with enhanced performance

ABSTRACT Whereas aluminum with combustion enthalpy of 32 KJ/g is the universal reactive metal particles, its passive protective oxide layer could encounter its efficient combustion. While titanium can offer low combustion enthalpy (20 KJ/g), it does not expose passive oxide layer. The impact of both reactive particles on ammonium perchlorate (AP) decomposition enthalpy was investigated via DSC. Aluminum and titanium particles secured an increase in AP decomposition enthalpy by 46.9%, and 80.7% respectively. This significant impact of titanium particles was ascribed to full exploitation of metal energy content. The impact of titanium particles on AP decomposition kinetics was assessed via Kissinger, KAS, and FWO models using TGA. Whereas virgin AP demonstrated average apparent activation energy of 155 KJ/mol, Ti/AP demonstrated average apparent activation energy of 141.2 KJ/mol. Solid rocket propellant formulations based on 17 wt % metal (Al or Ti) were developed via mixing and vacuum casting. Ballistic performance was investigated using small-scale ballistic evaluation rocket motor. Ti-based formulation offered an increase in propellant burning rate, average operating pressure, average thrust by 18.2%, 26.5%, and 19.7% respectively. It was concluded that titanium particles have a dual effect as efficient high energy dense material and as a catalyst.

Correction: Pang et al. Burning Rate Prediction of Solid Rocket Propellant (SRP) with High-Energy Materials Genome (HEMG). Crystals 2023, 13, 237

There was an error in the original publication [...]

Methodology for Studying Combustion of Solid Rocket Propellants using Artificial Neural Networks

The combustion properties of energetic materials have been extensively studied in the scientific literature. With the rapid advancement of data science and artificial intelligence techniques, predicting the performance of solid rocket propellants (SRPs) has become a key focus for researchers globally. Understanding and forecasting the characteristics of SRPs are crucial for analyzing and modeling combustion mechanisms, leading to the development of cutting-edge energetic materials. This study presents a methodology utilizing artificial neural networks (ANN) to create multifactor computational models (MCM) for predicting the burning rate of solid propellants. These models, based on existing burning rate data, can solve direct and inverse tasks, as well as conduct virtual experiments. The objective functions of the models focus on burning rate (direct tasks) and pressure (inverse tasks). This research lays the foundation for developing generalized combustion models to forecast the effects of various catalysts on a range of SRPs. Furthermore, this work represents a new direction in combustion science, contributing to the creation of a High-Energetic Materials Genome that accelerates the development of advanced propellants.