Solid Rocket Propellants Research Articles

Research progress on crack propagation behavior of solid propellant

Crack sprouting will lead to the mechanical properties of propellant degradation, and the additional combustion surface formed by crack propagation may cause over-pressurization, resulting in uneven combustion of propellant, significantly changing the thrust characteristics of the engine, reducing the range of the vehicle, and may even cause catastrophic accidents such as explosions. Therefore, the study of crack propagation law of solid rocket propellant pillar is an important basis to ensure the structural integrity and safety of the engine. In order to study the crack propagation pattern of propellant pillar under different working conditions, three aspects of propellant crack initiation criterion, crack propagation model, and crack propagation influencing factors are systematically reviewed. On this basis, the next research focus is sorted out, and it is considered that the construction of crack propagation model by establishing a multi-scale model, the characterization of crack propagation by using various test characterization means, and the analysis of damage evolution under the coupling of multiple working conditions will be the next research focus.

A review on viscoelastic behaviour of plasticizers in ap/al/htpb based composite solid propellant

Solid composite rocket propellant is commonly employed in military applications, such as in a rocket missile fuel engine. However, ongoing research is still being conducted on the material composition, fabrication process, and design of the grain structure to modify the performance of rocket motors. One of the most important parts of studying the solid rocket propellant is the propellant grain's fabrication process, which is the rheology and viscoelastic behaviour of the plasticizer as an additive material in the propellant slurry. Recent research indicates that the flowability of the propellant slurry has a substantial effect on grain quality. The plasticizer is an additive material that can improve the flowability of the propellant slurry by elongating the polymer chain structure. On top of that, most researchers have proven that the most crucial behaviour of the plasticizers is the particle migration that will affect the homogeneity of the propellant composition. This unwanted phenomenon was proven by the experimental work using chromatographic and Shore A hardness analysis of the propellant layer. The plasticizer concentration at 1 cm from the insulation layer decreased from 3 % to 2.087 % after the propellant was cured and aged. The Shore A hardness test shows the lowest value was at 3 cm from the insulation layer compared to 25 cm. Besides, it was also found that Isodecyl Pelargonate (IDP) can be a potential plasticizer for AP/AI/HTPB-based composite solid propellant application because it has the lowest viscosity value in the propellant slurry compared to the others. Thus, this paper focuses on reviewing the rheological behaviour of the plasticizers, influence of the curing temperature, plasticizer concentration, and curing time of the AP/AI/HTPB solid propellant slurry because all these factors affect the viscosity value of the plasticizers in the propellant slurry.

Парирование нештатных ситуаций при пусках исследовательских баллистических ракет с созданием резервных условий входа объекта испытаний в атмосферу

The article is devoted to solving the problem of determining and implementing backup conditions for the research ballistic missiles entry with the test objects in the inflight emergency situations. A test object is understood as a descent vehicle entering the atmosphere and moving at the speed close to the first cosmic velocity. In emergency situations during launch and on the active trajectory leg, backup entry conditions are making it possible to carry out the guaranteed flight testing, albeit with less efficiency, but with exception of the launch emergency completion and complete failure in fulfilling the flight test tasks. Detailed study of this problem is quite suitable for practical implementation and goes far beyond the scope of this work, since it requires design studies, including aerodynamic, dynamic, thermal and strength calculations, as well as the development of additional onboard and prelaunch algorithms. This paper notes only certain main features in decision making to terminate or continue the flight after the emergency situations for а research ballistic missile with liquid rocket engines and solid propellant rocket engines. Mathematical apparatus of the external ballistic design is used to solve the problem.

Burning Rate Prediction of Solid Rocket Propellant (SRP) with High-Energy Materials Genome (HEMG)

High-energy materials genome (HEMG) is an analytical and calculation tool that contains relationships between variables of the object, which allows researchers to calculate the values of one part of the variables through others, solve direct and inverse tasks, predict the characteristics of non-experimental objects, predict parameters to obtain an object with desired characteristics and execute virtual experiments for conditions which cannot be organized or have difficultly being organized. HEMG is based on experimental data on the burning rate of various high-energy materials (HEMs) under various conditions, on the metadata on the quantum and physicochemical characteristics of HEMs components as well as on thermodynamic characteristics of HEMs as a whole. The history and current status of the emergence of HEMG are presented herein. The fundamental basis of the artificial neural networks (ANN) as a methodological HEMG base, as well as some examples of HEMG conception used to create multifactor computational models (MCM) of solid rocket propellants (SRP) combustion, is presented.

Aluminum Particle Ignition Studies with Focus on Effect of Oxide Barrier

Aluminum particle ignition behavior in open atmosphere rocket propellants fires is of particular interest for preventing accidents for rockets carrying high-value payloads. For nominal motor pressures, aluminum particles oxidize to aluminum oxide in the gas phase and release significant combustion energy while minimizing motor instability. During rocket abort or launch pad malfunction which occur under atmospheric or low pressure, behavior of aluminum particle combustion becomes complex and aluminum appears to melt, agglomerate or form a skeletal structure. Furthermore, an oxide shell of alumina instantly forms on any fresh aluminum surface which is exposed to an oxidizing environment. Aluminum combustion then strongly depends on the oxide layer growth, which is influenced by causative factors, including particle size, environmental gas composition, and heating rate. This work focuses on the effect of the oxide barrier which forms on the surface of aluminum that is recognized to impede combustion of aluminum in solid rocket propellants. Understanding the mechanism for breach of this barrier is deemed to be an important consideration in the overall process. In this discussion, results of various experiments will be discussed which have a bearing on this process. Basically, a recognized criterion is the melting of the oxide layer at 2350 K is sufficient. However, in other situations, depending on the mechanism of oxide formation, there will occur defects in the oxide shell which provide for aluminum ignition at lower temperatures. For slow heating in an oxidizing environment, where the oxide layer can grow thick, then ignition is more difficult. Because there is no uniform model to establish an ignition criterion due to the unknown history of an aluminum particle, this paper reports experimental findings involving oxyacetylene torch, thermogravimetric analysis with differential scanning calorimeter, aluminum particle heating, electric ignition and aluminum powder heating, to address the influence of the oxide layer on the aluminum particle ignition.

Dual-functional organic surfactant for the effective stabilization of nano aluminium

An innovative approach that uses the dual functional capping agent 1-aminohexanol to stabilize aluminium nanoparticles (Nano-Al) produced via the colloidal synthesis process is reported.

COPING WITH SPACE ENVIRONMENT: TESTING SOLID PROPELLANTS FOR IN-ORBIT USE

The results of a risk reduction study are presented, proving the suitability of solid rocket motors (SRMs) to de-orbit satellites at the end of their lifetime. In a risk reduction program, three basic unexplored questions regarding this technological approach were investigated: (1) what are the effects of the size and type of particles ejected during the operation of SRMs; (2) what is the effect of the micrometeoroid impact on the solid propellant; and (3) what are the effects of prolonged exposure to space radiation during the 20-year expected satellite lifetime of satellites? Using a series of sophisticated and newly developed research approaches, it was shown that solid rocket propellants can withstand harsh space conditions over an extensive period of time. All of the tests were performed using a well-established, metal-free high-performance ammonium perchlorate/hydroxyl-terminated polybutadiene propellant fully qualified by military standards.

МЕТОДИКА ОПРЕДЕЛЕНИЯ ХАРАКТЕРИСТИК ВИБРОПРОЧНОСТИ КОНСТРУКЦИИ ПРИ ВЫСОКОСКОРОСТНЫХ ТРЕКОВЫХ ИСПЫТАНИЯХ АВИАЦИОННОЙ ТЕХНИКИ

Achieving the speed limits by ballistic aircraft is a priority, relevant and new technical challenge for the present. Track tests of such products are one of the final stages, confirming their practical performance and efficiency. Testing of products of aviation and rocket technology at the "Rocket Rail Track 3500" FKP "GkNIPAS" named after L.K. Safronov allow to accurately simulate the real conditions in which they are operated. The experimental setup includes a two-rail track, made on a special base, which excludes the unacceptable deflection of the R65 rail, with its fixing every 0.5 m by a special embedment design. The rail track has an acceleration section with an angle of attack of 2º and a length of 2500 m and a deceleration section. The test object is placed on the missile track carriage in such a way as to exclude the impact on it of shock waves reflected from the elements of the carriage and relief. Solid propellant rocket engines are installed on the carriage, which provide the necessary thrust. The sliding bearings of the rocket carriage cover the head of the rails. The thrust of the starting rocket engines provides the necessary acceleration to achieve the maximum values of the required test speed. Track high-speed tests of special equipment objects are accompanied by intense vibration and shock loading of the structure. As the speed of products increases above 600 m/s, as tests have shown, the amplitude of elastic oscillations of the structure can reach the maximum values permissible from the strength conditions. Experimental and theoretical study of vibration and shock effects on the structure of a track carriage with test objects under the conditions of an existing rail track is an urgent task for practice.

THERMAL PROTECTION SOLUTION FOR SOLID FUEL ROCKET ENGINES

This paper presents a new approach to the thermal protection of solid-fuel rocket engines. The method of thermal protection using coatings not only reduces the risk of damage to the engine casing but also increases the engine’s efficiency in working. Parameters such as thrust and the mass per square meter of the engine casing are used to compare the differences between engines with and without thermal protection coating. Numerical calculation applied to the cruise engine of the rocket 9M39. Calculation results show that the thickness of the thermal protective coating and the thickness of the combustion chamber have been optimized regarding rocket weight. The survey results showed that the mass per square meter of the engine casing decreased by 2.5 times, and the engine’s impulse total increased by 1.13% when using the thermal protective coating. These results are used as data to develop methods to improve the performance of solid propellant rocket engines.

Simplified stochastic temperature model for storage reliability estimation of solid rocket propellants

Simplified stochastic temperature model for storage reliability estimation of solid rocket propellants

Combustion process of a promising catalyst [Cu(Salen)] for HMX-CMDB propellants.

Metal organic complexes are regarded as a series of promising combustion catalysts for solid rocket propellants. Their effects on the combustion performance of propellants are closely related to the reaction mechanism. Here, the metal-organic complex Cu(Salen) was investigated as a candidate material for the combustion catalyst of the HMX-added composite modified double-base propellant (HMX-CMDB). The combustion performance of the propellant was found to be evidently enhanced in the presence of Cu(Salen) compared with the propellant samples containing Benzoic-Cu or without catalyst. The addition of Cu(Salen) can improve the burning rate and combustion efficiency of the propellant - and greatly reduce the burning rate pressure index. Analysis shows that the addition of Cu(Salen) can increase the combustion area, flame brightness and combustion surface uniformity of the propellant to a higher degree. The sample can spray more beams of bright filaments on the flat combustion section, and the amount of gas generated by decomposition also greatly increases. In addition, Cu(Salen) shows amazing advantages in improving the surface of the propellant and the temperature gradient of the combustion flame.

Design and optimization of a novel supersonic rocket with small caliber

<p style='text-indent:20px;'>It's difficult to meet the design requirements of a rocket with sensitive parameters using the traditional methods. This paper develops a design and optimization concept for a lightweight solid-propellant rocket with a small caliber and a high Mach number. The structural design and numerical modeling depend on the interior ballistic methodology. The model is validated by comparing pressure-time data from simulation and experiment. A modified evolutionary algorithm with constraints is applied because of the interactive design parameters and conflicting objects. The launch performance is improved by single- and multi-objective optimization. Under unchanged interior ballistic performance conditions, the peak pressure is reduced by 46.4%, and the erosive peak ratio is reduced by 46.4%. Despite the constraints, the Mach number is improved by 4.9%, the total impulse is improved by 6.3%, the peak pressure is reduced by 92.5%, and the erosive effects are reduced by 38.1% using different optimal solutions. A Pareto front is obtained by a constrained NSGA-Ⅱ, which reveals non-linear and non-uniform relations among design objects. A tidying method is proposed for a clear Pareto front. It indicates that, despite the sensitive parameters, launch safety and higher velocity are possible. The results help designers choose the best design schemes.</p>

Effects of Azide-Hydroxyl-Terminated Polybutadiene on the Properties of Solid Heterogeneous Rocket Propellants

Effects of Azide-Hydroxyl-Terminated Polybutadiene on the Properties of Solid Heterogeneous Rocket Propellants

Analytical Study of Sugar and Paraffin/Potassium Perchlorate as Solid Rocket Propellant

Nowadays, space exploration and sending satellites for different causes are more in demand. Current work is on Solid Rocket Propellant (SRP) which is used for boosters at the time of launch, correcting the satellite orbit with the help of thrusters and in missiles. Current work is to develop an easily available propellant, low cost, eco-friendly and simple to mold without compromising performance. An analytical study of the SRP parameters such as characteristic velocity, specific impulse, thrust at fixed O/F ratio is carried out. This paper proposed fuel as paraffin (C50H102) with sugar (C12H22O11) and oxidizer as potassium perchlorate (KCLO4). Thermal insulation is not used in the middle of the propellant and the outer casing because the heat does not sustain for long period for a small-scale rocket model. The specific impulse achieved with this SRP is 201.7 for an O/F ratio of 4.

INCREASING THE ENERGY OF alpha, omega - DIHYDROXYPOLIBUTADIENE (HTPB) USED IN SOLID HETEROGENEOUS ROCKET PROPELLANT

Solid heterogeneous rocket propellants (SHRP), which are very popular and widely used in the armaments industry (guided missile propulsion engines and long-range, short-range and medium-range anti-aircraft rockets) have inert binders that significantly affect the final performance parameters of propulsion charges. In this study, the popular HTPB binder was modified by introducing azide groups into the polymer chain during three different syntheses. Compounds with different content of explosive groups were obtained and tested for compatibility with the essential SHRP components: ammonium chlorate(VII) and dioctyl adipate. Then, preliminary application tests of the obtained HTPB derivative were carried out, showing the potential of the obtained energetic polymer and the possibility of its use as a binder in solid heterogeneous rocket fuels. HTPB, energetic binder, azide groups

Gas-dynamic losses in the flow part of the channel charge of a solid-propellant rocket engine

Modern methods of computational fluid dynamics have been used to study flows with mass supply in axisymmetric channels of solid-propellant rocket motor charges. The studies were carried out with the aim of increasing the accuracy of predicting the intraballistic parameters for performing engineering calculations. The analysis of changes in intraballistic characteristics in the flow part of the channel charge in the classical and nozzleless solid propellant rocket motors is carried out for different rates of mass supply from the combustion surface. For an isobaric combustion chamber, a characteristic velocity profile is shown along a tubular charge path and a charge with sudden expansion. It is shown that for a steady cosine profile of axial velocity after sudden expansion of the channel, a length of more than three gauges is required. For a high-speed combustion chamber, a comparison of the axial velocity profile is carried out depending on the flow velocity under different conditions of mass supply. The tendency of the influence of flow velocity on the character of the profile is noted. It is shown that at a Mach number over 0.5, an increase in the mass flow rate from the combustion surface provides a less filled velocity profile tending to cosine. The differences between the pressure losses in the flow passage of the channel charge, calculated in the axisymmetric approximation, and the losses determined using gas-dynamic functions, are shown. An increase of the mass flow rate in the charge channel of a nozzleless solid-propellant rocket motor leads to a decrease in pressure losses at Mach number over 0.8.

Combustion behavior and mechanism of molecular perovskite energetic material DAP-4-based composites with metal fuel Al

Combined with the oxidizer anions and fuel cations, molecular perovskite energetic materials show a good potential. In this work, the combustion behavior and mechanism of metal fuel aluminium (Al) with molecular perovskite energetic material (H2dabco)[NH4(ClO4)3] (DAP-4) as a high-energy oxidant was investigated. The DAP-4 based composites with metal fuel Al were designed and fabricated by the different mass ratios. Results showed that DAP-4 exhibits a good oxygen-supplied capacity for enhancing the combustion performance of Al. The maximum combustion heat of DAP-4/Al-3 at the Al/O mass ratio of 38:62 is up to 10,412 J/g in the inert gas, which is higher than those of other ratios and the mixtures of other energetic materials and Al. The evolution of pressure output, pressurization rate, and flame temperature was monitored for DAP-4/Al with different mass ratios. Composites DAP-4/Al/F were characterized by burning rates. The combustion reaction mechanism of metal fuel Al with DAP-4 as a high-energy oxidant was provided. DAP-4 was ignited firstly and released acid and oxidizing gases, which corroded Al2O3 shells on Al particle surfaces and accelerated the combustion reaction with Al to release a lot of energy. This work offered a new idea that molecular perovskite energetic materials have great potential in the high-energy Al-based solid rocket propellants.

Применение микрогибридных гетерогенных конденсированных систем для повышения энергетики ракетных двигателей

The paper considers heterogeneous condensed systems which potentially make it possible to use highly efficient oxidizers in their composition, including those liquid. The existing technologies are unable to include them in composition of the solid rocket propellants. Placing oxidizers such as nitrogen dioxide and tetranitromethane in hermetically sealed combustible polymer capsules connected by physical or chemical means allows creating a microhybrid rocket engine. Studies based on analyzing characteristics of similar systems and thermodynamic calculations demonstrate that such systems are potentially able to ensure mechanical strength required for the rocket engine functioning, sufficient speed, combustion static stability and chemical compatibility of components, as well as power at the level of liquid high-boiling fuel vapors, while maintaining advantages of the solid fuel engines. Considering their initial higher safety in production and operation, as well as the ability to be produced by modern technologies, microhybrid rocket engines appear to be promising area of further research.

Elaboration, Characterization and Thermal Decomposition Kinetics of New Nanoenergetic Composite Based on Hydrazine 3-Nitro-1,2,4-triazol-5-one and Nanostructured Cellulose Nitrate.

This research aims to develop new high-energy dense ordinary- and nano-energetic composites based on hydrazine 3-nitro-1,2,4-triazol-5-one (HNTO) and nitrated cellulose and nanostructured nitrocellulose (NC and NMCC). The elaborated energetic formulations (HNTO/NC and HNTO/NMCC) were fully characterized in terms of their chemical compatibility, morphology, thermal stability, and energetic performance. The experimental findings implied that the designed HNTO/NC and HNTO/NMCC formulations have good compatibilities with attractive characteristics such as density greater than 1.780 g/cm3 and impact sensitivity around 6 J. Furthermore, theoretical performance calculations (EXPLO5 V6.04) displayed that the optimal composition of the as-prepared energetic composites yielded excellent specific impulses and detonation velocities, which increased from 205.7 s and 7908 m/s for HNTO/NC to 209.6 s and 8064 m/s for HNTO/NMCC. Moreover, deep insight on the multi-step kinetic behaviors of the as-prepared formulations was provided based on the measured DSC data combined with isoconversional kinetic methods. It is revealed that both energetic composites undergo three consecutive exothermic events with satisfactory activation energies in the range of 139–166 kJ/mol for HNTO/NC and 119–134 kJ/mol for HNTO/NMCC. Overall, this research displayed that the new developed nanoenergetic composite based on nitrated cellulose nanostructure could serve as a promising candidate for practical applications in solid rocket propellants and composite explosives.

Application of Reichardt’s Solvent Polarity Scale (ET(30)) in the Selection of Bonding Agents for Composite Solid Rocket Propellants

Bonding agents (BA) are key compounding ingredients for the correct formulation of composite solid rocket propellants (CSRP). In particular, the addition of BA is essential to achieve suitable mechanical properties of CSRP in terms of adequate tensile strength and elongation at the break. It is shown that the polarity of each conventional BA as well as new potential BA can be measured through the Reichardt’s ET(30) polarity scale. Using this methodology, it was possible to propose a substitute for MAPO (tris-(methylaziridinyl)-phosphine oxide), a conventional BA with the drawback of high toxicity and high reactivity, with TTPT (tris-(pyrrolidine)-phosphine oxide), a completely safe and effective BA. In this work, several other potential BA were evaluated through the Reichardt’s ET(30) polarity scale but only a selection of the potential BA were effectively tested in a standard CSRP. The evaluation of TTPT vs. MAPO showing the ability of the former BA to match the mechanical properties of the latter BA was particularly interesting. A reasonable correlation between the elongation at break of the CSRP and the ET(30) value of the BA used in the compound was found.