Agglomerate Size Evolution in Solid Propellant Combustion under High Pressure

This work proposes a new effective method to realize variable thrust through discontinuous embedded metal wires in the solid rocket motor (SRM). We aimed to study the influence of discontinuous embedded metal wires on the performance of an SRM with a single-burning-rate grain. A model based on convection heat transfer, heat conduction, and heat radiation was established to calculate the heat transfer in the discontinuous embedded metal wires in the grain, to then obtain the burning rate ratio. Most importantly, a solid rocket motor was designed to verify the feasibility of variable thrust and of the present model prediction, with the embedded silver–nickel alloy wire divided into two segments in the grain. According to the SRM ignition experiment, the silver–nickel alloy wires raised the burning rate of the grain. The pressure varied regularly with changes in the discontinuous embedded metal wires. The theoretical burning rate ratio matched the experimental result well. Based on the verified model, the effects of the burning rate, pressure exponent, burning rate ratio, and number of wires on thrust were investigated. Burning rate, burning rate ratio, and pressure exponent were found to be positively correlated with thrust ratio. The thrust ratio could reach 12.5 when the burning rate ratio was 5. The ability to adjust thrust tended to increase with an increase in the number of wires. This study also provided a method to assess whether the consecutive embedded metal wires had been broken or not. The method using discontinuous embedded metal wires in the grain was proven to be feasible to realize multi-thrusts of single-burning-rate grain, which is a new idea for the design of a multi-thrust SRM.

Section: Introductionmentioning

confidence: 99%

Tuning the Ballistic Performance of a Single-Burning-Rate Grain Solid Rocket Motor via New Discontinuous Embedded Metal Wires

Wu,

Ren

2024

“…Their performance is closely related to the combustion characteristics of the propellants [1,2]. The combustion of propellants is a complex, multidimensional physical and chemical process at high temperatures and pressure in solid rocket motors [3,4], which makes it difficult to observe and measure the parameters by experiment. In such cases, the use of numerical simulation based on combustion models is an effective method to study the flame structure and combustion mechanism.…”

Section: Introductionmentioning

confidence: 99%

Numerical Investigation on the Effect of Ammonium Perchlorate Content and Position on the Combustion Characteristics of an Ammonium Perchlorate/Hydroxyl-Terminated Polybutadiene Propellant

Sun

Liu

et al. 2023

Self Cite

A gas–solid-coupled sandwich combustion model was established for ammonium perchlorate (AP)/hydroxyl-terminated polybutadiene (HTPB) composite propellant. Numerical simulations were conducted to investigate the influence of the content of AP and the relative position of the coarse AP on the flame structure and the burning rate of the propellant. The results indicated that the overall AP mass fraction has a significant effect on the gas-phase flame temperature and burning rate, and there exists an optimal oxygen-to-fuel ratio that maximizes the burning rate. As the mass fraction of fine AP increased, the premixed flame above the binder matrix gradually took over the dominance of the diffusion flame, and the intensity of the diffusion flame near the interface of coarse AP and binder matrix also increased, resulting in a significant increase in the burning rate. As the mass fraction of fine AP increases from 0% to 70.0%, the average surface temperature increases from 937 K to 1026 K, and the burning rate rises from 0.9 cm/s to 2.7 cm/s. The location of the coarse AP causes the flame tilts to the side with less binder matrix, but it had little effect on the burn rate of the propellant.

“…The burning rate and pressure exponent of a solid propellant are critical performance parameters for designing solid rocket motors [1]. The combustion mechanism of a solid propellant is complex and influenced by various factors, such as the formulation, pressure, initial temperature, gas flow velocity, and condensed combustion products [2,3]. However, despite extensive research, the current understanding of the underlying mechanisms is still limited.…”

Section: Introductionmentioning

confidence: 99%

Using the Impulse Method to Determine High-Pressure Dynamic Burning Rate of Solid Propellants

Liu,

Wang,

et al. 2023

A new method for determining the burning rate of a solid propellant, called the Impulse Method, is proposed in this paper. It is based on the proportional relationship between the impulse generated and the mass of the burned propellant. The pressure–time and thrust–time curves are obtained from a tubular propellant grain burning in the chamber, whose inner surface serves as the initial burning surface. Consequently, the mass of the propellant that was burned off at different pressures can be determined, and the burning rates at different pressures are derived according to the geometric parameters of the propellant grain. The Impulse Method was applied to test the burning rate of two types of propellants twice. The results show that the burning rates were consistent for the same propellant at corresponding pressures, demonstrating the feasibility and reliability of the Impulse Method. The burning rate of a GAP-based composite propellant at 20 MPa measured using the Standard Motor Method was 22.6 mm/s, and that measured using the Impulse Method was 22.2 mm/s and 22.7 mm/s, respectively. These findings indicate that the two methods have comparable accuracy. However, the Impulse Method has the advantage of obtaining the burning rate of the solid propellant at any pressure through a single test. In addition, the nozzle erosion only affected the pressure and not the burning rate. Finally, the rationality of the approach for determining the actual specific impulse was proven by comparing the results with those from another testing method.