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In this paper, the thermal decomposition characteristics and mechanical sensitivity of ETPE propellant were characterized by differential scanning calorimeter (DSC), vacuum stability tester (VST), impact sensitivity tester and friction sensitivity tester, and the main factors affecting the safety performance of ETPE propellant were studied. The results show that ETPE propellant has good thermal stability under 185°C. The combustion activation energy of ETPE propellant is slightly lower than that of HTPB propellant. By using 100°C heating method, the net added gas of ETPE propellant is 0.81ml/g, which meets the evaluation standard of stability qualification, and the propellant did not spontaneously combust under various high temperature test conditions. By reducing the amount of fine AP and controlling the amount of RDX, it is conducive to reducing the probability of the “hot spot” generated by the propellant under the external excitation energy, reducing the mechanical sensitivity of propellant and improving the safety performance of propellant.

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 nonlinear mechanical behavior and temperature sensitivity of HTPB propellant for solid rocket motors were investigated. The rate-dependent mechanical properties of the propellant were examined through a combination of experiments and numerical simulations. Experimental results demonstrate that the tensile mechanical properties of HTPB propellant are rate-dependent at 223 K and 323 K; stresses at a given strain gradually increase with increasing strain rate. By use a generalized nonlinear ZWT intrinsic model, the tensile mechanical behavior of HTPB propellant under a wide range of strain rates was to described. A numerical simulation of a uniaxial tensile test was performed using a UMAT subroutine. The results demonstrate that the model accurately represents the mechanical properties of the HTPB propellant.

In this paper, the HTPB propellant of a solid rocket motor was taken as the research object. Considering the effect of temperature on the propellant, various types of tests were carried out to study the damage law of the propellant at the macroscopic scale. The aging of HTPB propellant under temperature loading is mainly manifested as oxidative cross-linking of the binder network. The greater the cross-linking degree of the specimen is, the smaller the maximum elongation of the propellant is. Under constant strain load, some AP particles will be dehumidified, and at room temperature, AP particles will decompose slowly, resulting in an increase in strength and a decrease in elongation. The aging of HTPB propellant under constant strain loading at high temperatures is manifested as oxidation cross-linking of the binder network, ε m of propellant declines, and σ m of propellant is reduced.

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.

In order to improve the mechanical and combustion properties of composite solid propellant, the CuO@AP with core–shell structure was prepared by solvent–nonsolvent recrystallization method, and it was applied to AP/HTPB composite solid propellant. The thermal decomposition properties, sensitivity properties and tensile properties of CuO@AP propellant were studied and compared with ultrafine AP propellant, ultrafine spherical AP propellant and the mixture of CuO and AP (CuO/AP) propellant. The results show that the E a of ultrafine spherical AP propellant is 8.16% lower than that of ultrafine AP propellant with the same particle size, and the rate constant increases by 13.64%; the E a of the CuO@AP propellant is 23.63% lower than that of CuO/AP propellant with ultrafine AP of the same particle size, and the rate constant increases by 172.7%. What’s more, the catalytic effect of CuO@AP is obviously better than that of CuO/AP. The impact sensitivity of ultrafine spherical AP propellant is 29.61% lower than that of ultrafine AP propellant with the same particle size, and the ε b is increased by 51.35%. The impact sensitivity of the CuO@AP propellant is 25.38% lower than that of CuO/AP propellant with ultrafine AP of the same particle size, and the ε b is increased by 63.76%. The above shows that the CuO@AP composite particles with core–shell structure have potential application prospects in AP/HTPB propellant.

The effect of combustion catalysts which are several proportions of A-Pb, A-Cu and carbon on the combustion properties of composite modified double base (CMDB) propellant containing hexanitrohexaazaisowurtzitane (CL-20) were investigated by uniform design. The proportion of the three materials is optimized. The result shows that A-Pb, A-Cu and carbon can affect the combustion properties of CL-20-CMDB propellant. The effect of single A-Pb or carbon on the propellant combustion properties is obvious between 2 MPa and 8 MPa. The effect of single A-Cu on the propellant combustion properties is obvious between 14 MPa and 22 MPa. There is positive interactions of carbon with A-Pb or A-Cu between 2 MPa and 4 MPa and A-Pb with A-Cu or carbon are obvious. The calculation shows that the effect on adjusting the combustion properties of CL-20-CMDB propellant when the proportion of A-Pb, A-Cu and carbon is 3/0.8/0.2.

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 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.

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.