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This study investigates the influence of selected combustion rate catalysts on the ballistic, physicochemical, and mechanical properties of non-isocyanate heterogeneous solid rocket propellants. Methods for curing prepolymers and modifying hydroxyl-terminated polybutadiene (HTPB) to obtain carboxyl-terminated polybutadiene (CTPB) and its epoxidized derivative (EHTPB) are discussed. The initial stage involved the synthesis of CTPB and EHTPB. The obtained compounds were analyzed for viscosity, comparing their properties to those of the base polymer HTPB. FTIR spectra of the synthesized compounds were recorded. Crosslinking systems were formulated based on the synthesized substances and tested for tensile strength. The final stage consisted of preparing solid heterogeneous rocket propellants containing selected catalysts—catocene and iron nanopowder—and evaluating their burning rate, hardness, and density. The results of the rocket propellant tests indicate that both catalysts perform effectively in the proposed system. Significantly higher burning rates were achieved compared to the catalyst-free formulation. The addition of 1% catocene resulted in a 2.5-fold increase in burning rate. Even better performance was observed with iron nanopowder—1% addition led to an almost threefold increase in burning rate. Neither catalyst significantly affected the hardness of the propellant; all samples exhibited hardness values in the range of 71–76 Shore A. Increasing the catocene content led to a decrease in the final propellant density, whereas the addition of iron nanopowder increased the density relative to the base formulation.

This research paper presents the design and development of a solid rocket motor composed of sugar-based propellant for Rocket Assist Take Off (RATO) systems. RATO systems can be integrated with different platforms that require an initial boost to be launched such as the C130 aircraft. Sugar-based propellants offer a cost-effective and environmentally friendly alternative to conventional solid rocket propellants with effective launch performances. The research purpose is to design and develop prototype rockets, verify rocket performance, and optimize the design for upscaling to larger calibers. Additionally, to establish a standard method for manufacturing sugar-based propellant grains to ensure consistency and repeatability in every batch production. During the development, non-destructive testing was performed to check the grains’ quality and a series of static firings were conducted to measure the rocket performance. Three design concepts were developed for the 40 mm rocket to achieve sufficient thrust, enhance the overall performance, and achieve a boost-sustain profile. However, upscaling the motor to 75 mm posed new challenges in the design process, requiring modification to achieve the required thrust, ensure structural integrity, and improve grain quality. Further research is required to excel in large calibre rocket performance and characterize the propellant through BEM testing.

The effects of different types of nano-sized metal particles, such as aluminum (nAl), zirconium (nZr), titanium (nTi), and nickel (nNi), on the properties of a variety of solid rocket propellants (composite, fuel-rich, and composite modified double base (CMDB)) were analyzed and compared with those of propellants loaded with micro-sized Al (mAl) powder. Emphasis was placed on the investigation of burning rate, pressure exponent (n), and hazardous properties, which control whether a propellant can be adopted in solid rocket motors. It was found that nano-sized additives can affect the combustion behavior and increase the burning rate of propellants. Compared with the corresponding micro-sized ones, the nano-sized particles promote higher impact sensitivity and friction sensitivity. In this paper, 101 references are enclosed.

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 (AlH 3 ), 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.

Since the first synthesis of 2,6-diamino-3,5-dinitropyrazine-1-oxide (LLM-105) in Lawrence Livermore National Laboratory (LLNL) of America in 1993, efforts of investigation into LLM-105 have never ended due to its premium performance. Therefore, with the goal to further emphasize the importance of this material and to provide reference to the application of LLM-105 in propellants and explosives, the properties and performance of LLM-105 are briefly introduced at first and then the application of LLM-105 is reviewed in various areas, including polymer bonded explosives, double-base rocket propellants with low signature and low sensitivity, propellants for perforating cartridges in oil and gas fields and gun propellants.

In order to solve the problem of long time consuming and low efficiency in solving the complex combustion surface of solid rocket propellant, a combustion surface calculation method based on parameterized volume Boolean operation was proposed. After the solid model of propellant was created by using 3D modeling software, the parameterized burning surface model was used to perform Boolean difference calculation between the obtained burning surface growth entity and the propellant entity. Through Sensitivity Analysis function, the variation range of parameterized value was set, so as to obtain the burning face value under different burning thickness. In this paper, star shaped charge column is selected to calculate the combustion surface, and the feasibility of the method is verified. This method provides a basis for subsequent combustion surface calculation.

In this study, the feasibility of electrospraying as an alternative processing technique for the preparation of composite solid rocket propellants (SRPs) was investigated. The main objective was to improve microstructural homogeneity and interfacial contact between the oxidizer, energetic additive, and metallic fuel without altering the chemical composition of the formulation. Additionally, porous electrosprayed SRP formulations were prepared to examine the influence of controlled porosity on thermal decomposition behavior. The prepared materials were characterized using scanning electron microscopy combined with energy-dispersive X-ray spectroscopy (SEM/EDS) to assess microstructural features and component distribution. Thermal decomposition behavior and kinetic parameters were evaluated using simultaneous DSC/TG analysis conducted at multiple heating rates. Safety-related properties were assessed through friction sensitivity testing, while post-decomposition solid residues were analyzed using SEM/EDS and X-ray diffraction. The results show that electrospraying improves structural homogeneity, reduces solid residue formation after thermal decomposition, and decreases apparent activation energy, while maintaining unchanged friction sensitivity. These findings demonstrate the potential of electrospraying as a physical processing route for tailoring the microstructure and thermal behavior of composite solid rocket propellants.

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.

The effectiveness of multi-barrel rocket systems on today’s battlefields is strongly dependent on the reliability of operation and, hence, proper action of all components, especially rockets and propellants. Therefore, the properties of the solid rocket propellants used in the rocket motors must be determined with an efficient and reliable tool providing repeatable results. The article presents the results of a thermomechanical analysis of solid double-base rocket propellant used in multi-barrel rocket systems. One of the recommended methods for testing solid rocket propellants is dynamic mechanical analysis. Mechanical properties such as the dynamic storage modulus (E′), the dynamic loss modulus (E″), and the tangent tan(δ) of the phase shift angle (E″/E′) were measured with the use of the TA Instruments DMA Q 800 device, in a temperature range of − 100 to +100 ∘ C with the use of different frequencies of applied force and heating rates. Special attention was devoted to determining the glass transition temperature following the STANAG 4540 standardization agreement, as well as the influence of testing parameters on the obtained experimental results. Dynamic mechanical analysis has proven to be an effective method for the evaluation of key properties influencing rocket motor behavior.

Rockets have revolutionized space technology and human space exploration. Most rockets and missiles are both propelled by rocket motors that use composite solid propellants. The ICT code and the NSAS CEA code are two programs that can be used to forecast theoretical propulsion parameters for composite solid rocket propellant. Rocket propellant performance is governed by a specific impulse factor, which is calculated theoretical and experiment. In this paper, the theoretical specific impulse for different composite solid propellant formulations at 70 bar combustion pressure and an adapted nozzle (optimum expansion) were calculated by the NASA-CEA code and the ICT code. Meanwhile, a static firing test was performed on a small scale test motor to experimentally determine the actual specific impulse. The objective is to verify theoretical calculations from two codes with experimental data, via the determination of the specific impulse deviation co-efficient.