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The latest developments in solid propellants and their components are summarized. Particular attention is given to emerging energetic binders and novel, ‘green’ oxidizing agents and their use in propellant formulations. A brief overview of the latest reports on fuel additives is included. Finally, a summary of the state of the art and challenges in its development are speculated on.

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 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.

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.

Aluminium–lithium alloy (Al–Li alloy) powder has excellent ignition and combustion performance. The combustion product of Al–Li alloy powder combined with ammonium perchlorate is gaseous at the working temperature of solid rocket motors, which greatly reduces the loss of two-phase flow. Experimental investigations were thoroughly conducted to determine the effect of the Al–2.5Li (2.5 wt% lithium) content on propellant combustion and agglomeration based on thermogravimetry-differential scanning calorimetry, heat combustion, laser ignition, combustion diagnosis, a simulated 75 mm solid rocket motor and a condensed combustion products (CCPs) collection device. The results show that the exothermic heat and weight gain upon the thermal oxidation of Al–Li alloy is obviously higher than those of Al powder. Compared with the reference propellant’s formulation, Al–2.5Li leads to an increase in the burning rate and a decrease in the size of the condensed combustion products of the propellants. As the Al–2.5Li alloy content gradually increases from 0 wt% to 19 wt%, the burning rate increases from 5.391 ± 0.021 mm/s to 7.244 ± 0.052 mm/s at 7 MPa of pressure; meanwhile, the pressure exponent of the burning rate law is changed from 0.326 ± 0.047 to 0.483 ± 0.045, and the d43 of the combustion residue is reduced from 165.31 ± 36.18 μm to 12.95 ± 4.00 μm. Compared to the reference propellant’s formulation, the combustion efficiency of the HTPB propellant is increased by about 4.4% when the Al–2.5Li alloy content is increased from 0 to 19%. Therefore, Al–2.5Li alloy powder is a promising fuel for solid propellants.

Solid propellant grain, as a typical polymer, are the thrust generation devices and core load-bearing components of solid rocket motor (SRM) and are also known as SRM grain. They are constantly exposed to extreme service environments such as high temperatures, high pressures, and dynamic shocks, and have a relatively high failure rate in the field use of SRM. Its life and reliability are the shortcomings that restrict the improvement of weapons and equipment capability in China at present. This paper summarizes the typical fault types of SRM grain at present, and compares and analyzes the research progress of reliability design and analysis technology, reliability optimization technology, life test technology and reliability evaluation technology of SRM grain at home and abroad; This paper analyzes the deficiencies and reasons in the research and application of SRM grain reliability technology in China, and points out the technical difficulties and challenges faced by the integrated design of performance and reliability of SRM independent innovation design according to the needs of the forward research and development system of SRM. Based on the existing design level and industrial foundation in China, the basic research suggestions that should be carried out to consolidate the design ability of SRM grain in China are given.

The main objective of this paper is to compare the different methods used to measure the burning rate of hydroxyl-terminated polybutadiene (HTPB) propellant. The burning rate measurement is carried out under varying crosslinking density, which is determined by the curing ratio, i.e., the equivalent ratio of diisocyanate to total hydroxyl (NCO/OH) ratio. Four batches with different curing ratios were produced using fixed production processes. To measure the burning rate, cured solid propellant strands were tested using the acoustic wave emission method at different pressures ranging from 4 to 10 MPa and also by using static firing test of subscale motors. The study showed that there was little difference between the results obtained from the acoustic wave emission method and subscale motors.

Investigating the constitutive relationship and the damage failure mechanism of solid propellants is of significance for improving the safety, storage period and use efficiency of solid rocket motors. This paper focuses on the complex mechanical response behavior of composite solid propellants under loads and introduces experimental research on quasi-static and dynamic mechanical properties. Limited by the accuracy of instruments and testing methods, the research progress of macroscopic constitutive models, mesoscopic mechanical models and microscopic molecular models is summarized from the perspective of numerical simulations based on model scale and modeling methods. This paper tracks the historical progress of key models and summarizes the main achievements and prospects in this field. The research in this paper has high scientific and theoretical significance and engineering application value. It can provide an important reference and guidance for the structural optimization and performance improvement of solid propellants and lay a solid foundation for the development of solid rocket motors.

The employment of burning rate suppressants in the solid rocket propellant formulation is long known. Different research activities have been conducted to well understand the mechanism of suppression, but literature about the action of oxamide (OXA) and azodicarbonamide (ADA) on the thermal decomposition of composite propellant is still scarce. The focus of this study is on investigating the effect of burning rate suppressants on the thermal behavior and decomposition kinetics of composite solid propellants. Thermogravimetric analysis (TG) and differential thermal analysis have been used to identify the changes in the thermal and kinetic behaviors of coolant-based propellants. Two main decomposition stages were observed. It was found that OXA played an inhibition effect on both stages, whereas the ADA acts as a catalyst in the first stage and as coolant in the second one. The activation energy dependent on the conversion rate was estimated by two model-free integral methods: Kissinger–Akahira–Sunose (KAS) and Flynn–Wall–Ozawa (FWO) based on the TG data obtained at different heating rates. The mechanism of action of coolants on the decomposition of solid propellants was confirmed by the kinetic investigation as well.

A solid rocket motor is often in long-term storage, transportation, and other complex environment. Under the long-term action of these different states, the propellant in a solid rocket motor will face the risk of failure. One of the important reasons is the cumulative damage of solid propellant based on fatigue-creep interaction. This paper conducted a fatigue-creep interaction test on HTPB propellant under 0.7 MPa and 0.65 MPa stress, used a model describing the fatigue-creep interaction of metal materials to fit the test results under 0.7 MPa stress, and obtained a model describing fatigue-creep interaction of this type of HTPB propellant. The model was verified by using the test results under 0.65 MPa stress. The results show that there is a clear non-linear relationship as well as a positive correlation between the fatigue-creep interaction of the HTPB propellant, that is, the cumulative damage produced by the fatigue-creep interaction is more serious than the cumulative damage produced by a single fatigue or creep; the fatiguecreep interaction non-linear model used by the metal material has good applicability to the HTPB propellant.