Solid Rocket Propellants Research Articles

Influence of different crosslinking mixtures on the mechanical properties of composite solid rocket propellants based on HTPB

The crosslinking agent is a vital key which affects the mechanical properties of composite solid rocket propellants (CSRPs). Under this scheme, the effect of crosslinking mixtures (CMs) based on trimethylolpropane (TMP) as a triol crosslinker and butanediol (BD) as a chain extender on CSRPs based on hydroxyl-terminated polybutadiene was investigated. A series of 27 propellant compositions was formulated to study the mechanical properties of the prepared CSRPs. The effect of changing the weight ratio of TMP to BD in the CM was studied. In addition, the influence of increasing the percentage of CM (from 0% to 0.5%) in the prepared samples was investigated. Also, the effect of the CM on CSRPs containing different curing ratio of NCO/OH = 0.7, 0.75, and 0.8 was studied to generate the largest possible strain-ability with high strength over different levels of curing conditions. The mechanical characteristics (tensile strength and strain) of the prepared CSRPs have been measured and plotted versus CM content, NCO/OH and TMP:BD ratio. Generally, the addition of CM leads to a remarkable enhancement in the propellant mechanical properties. Samples containing TMP:BD (2:1) provide the highest strength while samples containing TMP:BD (1:2) show the highest strain over all the NCO/OH ratios. Formulations with TMP:BD (1:1) give high strength with moderate strain. Variation in CM content has a remarkable influence on the mechanical properties of CSRPs. A wide range of tensile strength and strain were obtained from this study to offer variety of results suitable for different applications in the CSRPs technology.

EMERGENCY BURNING OF SOLID ROCKET PROPELLANT: DAMAGE RISK ASSESSMENT TO PEOPLE IN THE WORKPLACE

Purpose. This work includes the development of a computer model to calculate the risk of thermal damage to people in the shop in case of emergency burning of solid rocket propellant. Methodology. To calculate the temperature field in the shop in order to determine the zones of thermal damage to workers in the building, the equation expressing the law of energy conservation was used. Based on this modeling equation, the temperature field in the shop is calculated in the presence of a source of heat emission – burning solid rocket propellant. To calculate the velocity field of air flow in the shop, taking into account the location of obstacles in the path of heat wave propagation, we used the model of vortex-free air motion – the equation of the velocity potential. A two-step finite difference scheme of conditional approximation is used to numerically solve the equation for the velocity potential. A difference splitting scheme was used to numerically solve the energy equation. At the first stage of construction of the difference splitting scheme of the two-dimensional energy equation into the system of one-dimensional equations is performed. Each one-dimensional equation allows you to calculate the temperature change in one coordinate direction. The point-to-point computation scheme is used to determine the temperature. When conducting a computational experiment, the air exchange in the building is taken into account. The risk assessment of thermal damage to personnel in the building is performed for different probabilities of the place of emergency combustion of solid rocket propellant. Findings. Using numerical model prediction of the potential risk areas of thermal damage to staff in the shop for a variety of emergency situations was performed. Originality. A computer model for rapid assessment of the potential risk of damage to people in the shop in case of emergency burning of solid rocket propellant was constructed. Practical value. The authors developed a code that allows you to quickly simulate the temperature fields formation in the shop in case of emergency burning of solid rocket propellant and to identify potential areas of thermal damages to workers based on this information. The developed computer program can be used to assess the risk of thermal damage in the chemical industry in case of emergency.

Enhancing combustion performance of nano-Al/PVDF composites with β-PVDF

Solid energetic nanocomposites, consisting of reactive metal particles and oxidizers, find broad applications ranging from pyrotechnics to solid rocket propellants and solid ramjet fuels. In particular, nano-aluminum (n-Al) and polyvinylidene fluoride (PVDF) are attractive fuel and oxidizer materials, due to the high energy density of n-Al, and the high oxidizing potential and excellent mechanical properties of PVDF. PVDF is a semi-crystalline polymer and has three common crystalline phases, alpha (α), gamma (γ), and beta (β) phases. Many research works have focused on the chemical reactions between Al and PVDF, yet the effect of PVDF crystallinity on the Al/PVDF reaction is unknown. Here, we experimentally and computationally demonstrate that increasing the mass fraction of β-phase PVDF from 2.5 to 25% in Al/PVDF composites substantially improves peak pressure by 90% (from ~34 to ~64 psi) and pressure rise rates by 300% (from 2.5 psi/ms to 10 psi/ms). This stems from the alignment of F atoms along one side of the β-PVDF polymer chain, making it structurally conducive to reacting with Al particles to form strong Al-F interactions. This strong interaction leads to higher binding energy between, and hence higher reactivity in, β-PVDF and Al. Our research provides a new method for enhancing the reactive performance of Al/PVDF composites by increasing the content of β-PVDF.

Combustion and catalytic performance of metal-free heat-resistant energetic polymeric materials

Energetic polymeric materials play vital roles in energetic materials fields to improve the comprehensive performances of rocket propellants and PBXs. Herein, four N-di/trinitrobenzene based heat-resistance energetic polymeric materials were synthesized via two facile approaches and fully characterized using FTIR, NMR, SEM and XRD. The combustion properties of energetic polymers were tested and recorded on a high-speed camera, physicochemical and energetic properties were also investigated. The results show that as-synthesized energetic polymers not only exhibit excellent thermal stabilities (the highest exothermic decomposition temperature of 336.7 °C) superior to those of GAP, poly(GLYN), triaminoguanidine-glyoxal polymer (TAGP) and graphene oxide (GO), but also produce great detonation and compatibility performance. In addition, the energetic polymer samples can be directly ignited without any additives. Worth noting that, we reported the thermal decomposition catalysis of AP by non-metallic energetic polymers as the first example, suggesting they can be potentially developed into metal-free catalysts replacing heavy metal-based energetic coordination polymers (ECPs) or energetic metal-organic frameworks (EMOFs). In summary, as-prepared energetic polymeric materials represent great application potential in solid rocket propellants and PBXs due to their promising catalytic effect, excellent chemical and physical performances.

Impulse and Performance Measurements of Electric Solid Propellant in a Laboratory Electrothermal Ablation-Fed Pulsed Plasma Thruster

Electric solid propellants are advanced solid chemical rocket propellants that can be controlled (ignited, throttled and extinguished) through the application and removal of an electric current. This behavior may enable the propellant to be used in multimode propulsion systems utilizing the ablative pulsed plasma thruster. The performance of an electric solid propellant operating in an electrothermal ablation-fed pulsed plasma thruster was investigated using an inverted pendulum micro-newton thrust stand. The impulse bit and specific impulse of the device using the electric solid propellant were measured for short-duration test runs of 100 pulses and longer-duration runs to end-of-life, at energy levels of 5, 10, 15 and 20 J. Also, the device was operated using the current state-of-the-art ablation-fed pulsed plasma thruster propellant, polytetrafluoroethylene (PTFE). Impulse bit measurements for PTFE indicate 100 ± 20 µN-s at an initial energy level of 5 J, which increases linearly with energy by approximately 30 µN-s/J. Within the error of the experiment, measurements of the impulse bit for the electric solid propellant are identical to PTFE. Specific impulse when operating on PTFE is calculated to be about 450 s. It is demonstrated that a surface layer in the hygroscopic electric solid propellant is rapidly ablated over the first few discharges of the device, which decreases the average specific impulse relative to the traditional polytetrafluoroethylene propellant. Correcting these data by subtracting the early discharge ablation mass loss measurements yields a corrected electric solid propellant specific impulse of approximately 300 s.

Effect of Complex Metal Hydrides on the Elimination of Hydrochloric Acid Exhaust Products from High‐Performance Composite Solid Propellants: A Theoretical Analysis

AbstractAmmonium perchlorate (AP)‐based solid rocket propellants generate hydrochloric acid (HCl) as an exhaust product during combustion. This latter displays numerous environmental problems such as the depletion of the ozone layer and the increase of the concentration of acid rains. This paper offers a theoretical analysis concerning the employment of metal hydrides as efficient additives to mitigate some of the negative effects of HCl on the environment during the combustion of AP‐based composite propellants. Nine complex metal hydrides, expected to generate desirable performance gains, are selected to assess their scavenging effect on the HCl during propellant combustion. Based on NASA Lewis Code, Chemical Equilibrium with Application (CEA), comparative analysis of the theoretical performance of specific impulse, adiabatic flame temperatures, condensed combustion products as well as the exhaust gaseous species has been carried out. This study reveals that complex metal hydrides‐based propellants exhibited better performance than those containing simple metal hydrides, showing a synergetic effect of high specific impulse and low environmental impact. It is consequently anticipated that a good choice of metal hydride can provide clean/green propellant formulations that can satisfy the current environmental requirements with better performance than current aluminum propellants.

Decomposition of Ammonium-Perchlorate-Encapsulated Nanoscale and Micron-Scale Catalyst Particles

Iron oxide is the most common catalyst in solid rocket propellant. Increased performance of propellant by encapsulating iron oxide particles within ammonium perchlorate (AP) has been previously demonstrated, but only nanoscale particles were used, and encapsulation was only accomplished in fine AP ( microns in diameter). In this study, the size of the particle inclusions were extended to micron scale within the AP particles, and the particle sizes of the AP-encapsulated catalyst particles were increased to the coarse region (hundreds of microns). Fractional crystallization techniques were used with fine AP-encapsulated particles acting as nucleation sites for precipitation. This paper reports catalyst particle inclusions of micron scale, as well as nanoscale, within AP and presents characterization of this encapsulation. Encapsulating micron-sized particles and growing these composite particles could pave the way for numerous possible applications. A study of the thermal degradation of these AP-encapsulated particles compared against a standard mixture of iron oxide and AP showed that AP-encapsulated micron-scale catalyst particles exhibited similar behavior to AP-encapsulated nanoscale particles. Using computed tomography, it was found that catalyst particles were dispersed throughout the interior of coarse AP-encapsulated micron-scale catalyst particles and decomposition was induced within these particles around catalyst-rich regions.

Synthesis of multiphase binary eutectic Al–Mg alloy-nanoparticles by electrical wire explosion technique for high-energy applications, its characterisation and size-dependent thermodynamic and kinetic study

Abstract In the present study, Al–Mg binary eutectic alloy nanoparticles were synthesized through the electrical wire explosion technique (EWET) in an inert ambience. High-speed imaging was carried out to visualize sublimation process of alloy wire and subsequent formation of nanoparticle through the condensation process. The fundamental gas-phase kinetics is investigated by the embedded atom method (EAM) and reactive hard-sphere model to elucidate the mechanisms governing the formation of finer particles of eutectic Al-Mg alloy. The condensed phase transformations are analysed through size-dependent thermodynamics and growth kinetics modelling, which has shown that the adsorption controlled growth is responsible for the generation of polydispersed particles. The formation of Al20Mg23 nanoparticles and its morphological features were characterised through XRD and SEM/EDS analysis. High-resolution transmission electron microscopy (HRTEM) along with the selected area electron diffraction (SAED) confirmed the spherical morphology, crystallinity, lognormal particle size distribution and the multiphase microstructure of the eutectic AlMg nanoparticles. The formation of multiphase alloy nanoparticles is attributed to the insignificant difference in melting temperatures of Al and Mg and phase boundary modifications due to size-effect. Scanning transmission electron microscopy (STEM) revealed the spatial distribution of Al and Mg within the alloy nanoparticle. Differential scanning calorimetry (DSC) confirmed the reduction in melting temperature and fusion enthalpy of eutectic Al–Mg alloy nanoparticles. It is proposed as a potential substitute for Al nanoparticles to reduce the ignition temperature, agglomeration and two-phase flow losses for realising improved combustion performance in solid rocket propellants.

Synthesis and combustion catalytic activity of ferrocene‐based energetic compounds

AbstractAmmonium perchlorate (AP) is a common oxidizer in composite solid rocket propellants due to its excellent burning characteristics, good processability, and storability. Owing to their outstanding catalytic effects, ferrocene, and its derivatives have become the most widely used burning rate catalysts (BRCs). The addition of ferrocene and its derivatives to AP rendered performance optimization. In this study, azole‐based ferrocenyl compounds were successfully synthesized. The compounds were characterized by single‐crystal X‐ray diffraction, UV‐vis spectroscopy, and other techniques. The thermal degradation of AP catalyzed by these compounds was evaluated by differential scanning calorimetry and thermogravimetric analysis. Results revealed that the decomposition peak temperature of AP dramatically decreases and that the released heat of AP significantly increases with the new compounds as additives. Hence, the six azole‐based ferrocenyl BR catalysts are favorable for the combustion catalytic activity.

Investigation of Solid Propellant Rocket Motor Nozzle via CFD Simulation

The current work predicts the combustion characteristics of a solid rocket motor especially the flow in the nozzle part. The simulation work is done using computational fluids dynamic (CFD) software. The volume mesh was built in one quarter of the nozzle to reduce time for calculation. The same nozzle is also created with different mesh sizes for grid dependency check. All the rocket parameter is obtained from the theoretical design of rocket motor based on basic nozzle design. The simulation result in CFD is compared to the theoretical calculations and it shows that the modelling of rocket nozzle results shows good agreement with the theoretical ones. Parametric study on the nozzle was conducted. The results show that the best nozzle configuration for the current design is when the throat length is set to 1mm.

Core-shell nAl@Fc-Fx nanocomposites with dual function: Combustion and anti-migration performance

In this study, three fluoro-containing ferrocene compounds (Fc-Fx) with different fluorine and Fc contents were prepared and then be coated on the surface of nanoaluminum (nAl) to increase the combustion performance of nAl and the anti-migration performance of Fc. By removing the aluminium oxide (Al2O3) shell and instead constructing a new Fc-Fx shell with a thickness of 5–30 nm on the surface of nAl, the obtained nAl@Fc-Fx nanocomposites showed enhanced performances in i) anti-migration and ii) combustion performances owing to the Al-F interaction and the intimate contact between nAl and Fc-Fx. Among all three nAl@Fc-Fx nanocomposites, nAl@Fc-F2 with the longest fluorine chain could shorten the ignition delay time from 1.83 s to 1.03 s and nAl@Fc-F3 with highest Fc content could decrease the decomposition temperature of ammonium perchlorate to 328 °C. When the nAl@Fc-Fx nanocomposites were applied in the propellants, a shorter migration length (0.1–0.4 cm) was achieved as compared with the physically mixing nAl/Fc compound (3.0 cm), which was beneficial to enhancing their burning performance. This work demonstrated that nAl@Fc-Fx nanocomposites had great potential to be applied in solid rocket propellants.

Experimental Studies of the Scatter in Heat-Shielding Characteristics of a Rubber-Like Heat Shield

The scatter in heat-shielding characteristics of a material used as a heat shield of high-temperature units, in particular, solid propellant rockets, considerably affects the design thickness and weight of the heat shield, and, hence, the efficiency and reliability of the unit. Based on the results obtained during high-temperature bench tests, the data scatter is shown to depend considerably on the method used to calculate the depth of heating and/or the depth of thermal destruction. Analysis of the experimental data yielded estimates of the scatter in the heat-shielding characteristics of a rubber-like thermal-protection material.

Modelling and experiments of enclosure solid-propellant fires

In order to simulate an accidental propellant fire, blocks of aluminized solid rocket propellant, with mass in the range of 2-11 kg, were burned in a full-scale enclosure representative of an ammunition storage room onboard naval vessel. Wall heat flux and temperature, as well as gas pressure and temperature were measured to provide information regarding fire behaviour and thermal response of the structure. Burn characteristics of the propellant block, including shape and geometry, were deduced from visible and infrared images. The role of confinement, interior wall insulation or water spray on fire consequences was also investigated. Simulation results from a global model were compared to some experimental data. Agreement and discrepancies were identified.

Exploring the telechelic block copolymers of polybutadiene and ϵ-caprolactone for composite propellant application

In this study, synthesis of three hydroxyl-terminated block copolymers of varying molecular weight by ring opening polymerization of ԑ-caprolactone using hydroxyl-terminated poly butadiene as the initiator and stannous octoate as the catalyst was attempted. These block copolymers were specifically synthesized for its application in solid rocket propellants as an energetic binder. The desired chemical links between blocks were noticed in 1H nuclear magnetic resonance (NMR) and 13C NMR spectra, followed by confirmation in 1H/13C heteronuclear single quantum coherence and Fourier transform infrared spectra. To specify these block copolymers, various properties such as hydroxyl value (OH value), number average molecular weight (M¯n), specific gravity (ρ), rheological properties, glass transition temperature (Tg), thermal decomposition, and miscibility with nitrate ester plasticizers were determined. Additionally, the mechanical properties of the samples were evaluated using isocyanates as curators at NCO/OH equivalence ratio of 0.8 and 1.0. Finally, the best combination was selected based on physicochemical properties and miscibility with nitrate esters for further evaluation in composite propellant. We also theoretically studied the ballistic parameters of composite propellants by the NASA CEC-71 program to observe the gain of density impulse as the plasticizer to polymer ratio was increased.

Determining non-stationary state of solid-propellant rocket engine model based on numerical conjugate problem solution

This paper describes the solid-propellant rocket engine structure, including casing and filler. The gas flows within the filler channel. Based on numerical conjugate problem solution, non-stationary stress-strain states of the casing and filler were determined, as well as gas flow parameters in the channel. Within the conjugate problem, the engine is considered to be a two-component system: deformable solid body and gas. Conjugate problem solution involves specific subtasks related to the conjugation conditions, where Lagrangian approach is used for subtasks of solid body. Euler approach is used for subtasks of gas, implying immovable boundary of the computational domain on the integration time step. Numerical methods are applied in solving subtasks. Specific feature of conjugate problem solution algorithm is discrete movable domain boundary interface.

Unusual Energetic Periodate, Sulfate and Amino‐bistetrazolate Salts of the Trinitropropylammonium Cation

Starting from a nucleophilic Michael addition of nitroform to acrylamide, three synthetic strategies towards 1,1,1‐trinitropropyl‐1‐ammonium salts are described. Due to the high oxygen content and density of one periodate salt, its suitability as high‐energy dense oxidizer in energetic formulations was predicted according to its specific impulse and detonation velocity. Furthermore, those properties were compared to the nitrogen‐rich amino‐bis(5‐tetrazolate) salt. The parameters were calculated with the EXPLO5 (V6.03) computer code and compared to the common solid rocket propellant ammonium perchlorate (AP). Calculations towards the energies of formation were performed using Gaussian 09. Characterization including multinuclear NMR spectroscopy, vibrational analysis (IR, Raman) as well as elemental analysis and a single‐crystal X‐ray diffraction study was performed. The thermal stability was studied using differential scanning calorimetry and the sensitivities against impact and friction were determined.

Thermodynamic Properties of Hydroxylammonium Nitrate-Based Electric Solid Propellant Plasma

Electric solid propellants are advanced solid chemical rocket propellants controlled by electric current. An electric solid propellant may also be used in an electric propulsion system: specifically, an ablative pulsed plasma thruster. Previous experiments with the electric solid propellant known as high-performance electric propellant, or HIPEP, suggest its ablation processes are similar to the traditional propellant polytetrafluoroethylene (). Better understanding of the ablation and the resulting propulsion performance of HIPEP requires a model of its vapor composition, bulk plasma quantities, and thermodynamic properties. Such a model was developed and favorably compared with previous analogous model predictions found in the literature for . The electric solid propellant vapor composition was predicted in the temperature range of 500–40,000 K at 1 bar pressure. Low temperatures (less than 2000 K) were dominated by water (), carbon dioxide (), and molecular nitrogen (), with results at 700 K matching to within 10% of the previous combustion model predictions. At high temperatures (greater than 25,000 K), the vapor was strongly ionized and dominated by carbon (), oxygen (), nitrogen (), and hydrogen () ions. The calculated enthalpy and specific heat were higher for high performance electric propellant (HIPEP) vapor than , suggesting increased thermal losses could be inherent to ablation-controlled HIPEP discharges.

The transfer and exploitation of German air-to-air rocket and guided missile technology by the Western Allies after World War II

During World War II, the German military-industrial complex under the national socialist regime made pioneering strides in the development of aircraft armaments for the Luftwaffe. Through the application of scientific knowledge in the fields of mechanical and electrical engineering, ballistics, aerodynamics, and rocket propulsion, German armaments manufacturers evolved aircraft cannon technology to the automatic revolving aircraft cannon, to small unguided solid propellant rockets, and by 1945 Germany was on the threshold of starting series production of the world’s first guided air-to-air missile system. Towards the conclusion of the war in Europe and in the months afterwards, the four occupying powers produced scientific and technical intelligence on these developments for the purpose of exploitation. This article provides previously unpublished details about these activities by the Western Allies – the United States, United Kingdom and France – historical events which have traditionally been overshadowed by the Allies’ transfer and exploitation of other German guided weapons, such as the V-1 and the V-2.

ЧИСЛЕННОЕ ИССЛЕДОВАНИЕ НЕСТАЦИОНАРНЫХ ГАЗОДИНАМИЧЕСКИХ ПРОЦЕССОВ ПРИ СТАРТЕ ТВЕРДОТОПЛИВНЫХ РАКЕТ

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.

Numerical simulation of the unsteady combustion of solid rocket propellants at a harmonic pressure change

This study focuses on a numerical investigation of the unsteady burning rate of solid propellants at a harmonic pressure change in the combustion chamber of a solid propellant rocket engine. The physico-mathematical model includes the equations of heat transfer and decomposition of the oxidizer in the solid phase and two phases, the dual velocity, and the two-temperature reaction flow of gasification products. The boundary conditions on the solid fuel surface implement the conservation of energy fluxes and the mass of components. We numerically calculate the unsteady burning rate of metallized solid propellant and nitroglycerin powder under a harmonic pressure change in the combustion chamber of a solid propellant rocket engine and determine the dependence of the burning rate amplitude on the frequency of pressure oscillations. The amplitude of the burning rate depends nonmonotonously on the oscillation frequency. With increasing frequency, the amplitude first rises and then declines.