Properties Of Propellant Research Articles

Uniaxial tensile mechanical properties of NEPE propellant under wide temperature and wide strain rate

Abstract In order to gain insight into the mechanical properties of the NEPE propellant under a range of temperature and strain rate conditions, the uniaxial tensile mechanical property test of NEPE propellant with a wide temperature range of −40°C-70°C and a wide strain rate range of 0.0012s−1-1s−1 has been carried out based on the in-situ testing device of solid propellant microfabrication under complex working conditions and the supporting high and low temperature integrated machine. Additionally, the morphology of the solid particles of the propellant was observed and analyzed during stretching using the microscope equipment provided by the stretching machine. The results show that it has been demonstrated that both temperature and strain rate exert a more pronounced influence on the mechanical properties of propellants; at low temperatures, it has been demonstrated that the stress-strain curve of the propellant exhibits a more pronounced inflection point as the strain rate increases; the maximum tensile strength of the propellant demonstrates a gradual increase with a reduction in temperature and an increase in strain rate. Furthermore, a linear double logarithmic relationship is observed with the strain rate. In contrast, the maximum elongation exhibits no discernible pattern of change with temperature and strain rate; when observed during stretching, it was found that the main forms of propellant damage destruction at low temperatures were dehumidification between the particles and the matrix and particle rupture.

Research on mechanical properties of HTPB propellants

Abstract Slat propellants with different sizes and shapes are designed to investigate the strength criteria of HTPB propellants. According to the calculation and simulation results, with the increase of multiple, the designed size slat specimens can achieve equivalent biaxial tensile stress ratios of 1:2, 1:2.5, 1:3, 1:3.5, and 1:4 under uniaxial tensile conditions. The biaxial mechanical properties of the slab with a stress ratio of 1:2 were tested at low temperatures (25°C, −30°C, −50°C, and 0.4 s−1, 4.0 s−1, 14.29 s−1, 42.86 s−1). It was found that as the temperature and the strain rate rise, an increase is observed in the maximum tensile strength and the initial modulus, a decrease is observed in the maximum elongation of the propellant, and the strength envelope is constructed according to the typical mechanical parameters: the strength envelope of HTPB propellant increases as the temperature decreases. The research findings are expected to provide statistical support for the analysis of the structural integrity of HTPB propellants under low-temperature ignition.

Unveiling the impact of nitrocellulose-enveloped nCuO on the thermal characteristics of ammonium nitrate-based solid composite propellants utilizing polyurethane/nitrocellulose blend binders

ABSTRACT This study investigates the combined effects of coated nano copper oxide (NC@nCuO) and polyurethane (PU) modification on the thermal degradation kinetics and combustion properties of ammonium nitrate-based solid composite propellant. Nitrocellulose (NC) is used as a coating for nCuO, forming a core-shell structure, and as a binder in a PU/NC blend. Surface modification of ammonium nitrate was achieved through repeated spray coating. The formulations underwent thorough thermal characterization using thermogravimetry (TG) and differential scanning calorimetry (DSC) analyses. TG analysis showed a significant decrease in thermal degradation temperature due to AN modification with NC@nCuO and the use of PU/NC as a binder. Isoconversional kinetic methods (it-KAS, it-FWO, and VYA/CE) based on DSC tests were used to predict kinetic parameters, revealing a substantial impact of NC-coated nCuO and PU modification on the reactivity of the propellant, reducing the activation energy from 125 kJ/mol (reference) to 99 kJ/mol (nCuO) and 94 kJ/mol (NC@nCuO). Additionally, combustion tests were conducted. The results showed that the introduction of nCuO reduces the ignition delay by 5.25 ms and the combustion duration by 0.06 ms, while NC@nCuO further reduces these by 8.17 ms and 0.12 ms, respectively. Additionally, there is a 23% and 65% increase in flame intensity relative to the reference formulation after the incorporation of nCuO and NC@nCuO, repectively.

Theoretical and experimental study on combustion response of aluminized solid propellant under acceleration effects

The combustion performance of solid propellants is significantly affected by the external operating parameters of rocket motors. In this paper, the combustion response characteristics of aluminized solid propellant under acceleration conditions have been studied theoretically and experimentally. A theoretical model of unsteady burning for aluminized solid propellant, considering transient heat transfer and acceleration effects, is developed based on the Zel'dovich-Novozhilov solid-phase energy conservation. The pressure-coupled responses of the solid propellant are also measured experimentally under different normal accelerations using the T-burner and centrifuge. The model is well validated by comparison with literature and experimental results. The unsteady burning of propellants is analyzed in detail, and the effects of acceleration and propellant properties on combustion response are investigated. The results show that there is a significant difference between the transient and quasi-steady burning rate under dynamic environments. As normal acceleration increases from 0 g to 1000 g, the fluctuation amplitude of the net heat flux on the burning surface continues to decrease, the magnitude of the pressure-coupled response peak decreases significantly by 69 %, and the frequency of the response peak shifts from 140 Hz to 230 Hz. Moreover, the pressure exponent and thermal conductivity plays a positive role in the pressure-coupled response. The model presented in this paper can be helpful for predicting the differences between flight and ground test performance and evaluating the combustion stability of solid rocket motors.

Damage analysis of solid propellants with default defects based on macro-microscopic approach

Solid Rocket Motors (SRMs) rely on solid propellants, directly impacting the energy capacity of launch vehicles. High energy density material (HEDM) propellant exhibits superior energy density owing to the incorporation of a substantial quantity of energetic particles. However, the elevated particle fill ratio makes the formation of pores more prevalent during the manufacturing process of propellants. These pores, representing default defects, are indispensable factors in the investigation of mechanical properties of propellants. This study investigated default defects effects on HEDM propellants, analyzing both microscopic and macroscopic perspectives. Utilizing micro-CT imaging, the microstructure and default defects were characterized. Subsequently, using the cohesive zone model (CZM), microscopic damage evolution was depicted with the representative volume element (RVE) model, allowing in-depth analysis. Furthermore, a damage constitutive model, based on microscopic damage, was developed, with its softening function calibrated using microscopic damage evolution characteristics. Implemented into ABAQUS via user-defined material subroutine (UMAT), this model enabled the prediction of the mechanical response of HEDM propellants with different defects. The results demonstrate that the developed model accurately predicts the mechanical responses of HEDM propellants with different defects, showcasing the potential of macro-microscopic approaches in analyzing propellants with default defects.

Structural integrity assessment of a solid propellant grain considering confining pressure effect

Confining pressure has a significant effect on the mechanical properties of solid propellant, and it is crucial to develop a refined structural integrity analysis and assessment method considering confining pressure effect and nonlinearity of propellant material for the design of propellant grain of a solid rocket motor (SRM), especially for the design of high charge modulus. In this work, an existing viscoelastic-viscodamage constitutive model considering confining pressure effect was implemented as a user-defined material subroutine and verified using a few experiments under various confining pressures. The mechanical responses and safety factor of a typical variable cross-section propellant charge SRM with various charge modulus under ignition pressurization load and room temperature were analyzed and evaluated. The results show that the user-defined material subroutine can well describe nonlinear mechanical responses of solid propellant under different experimental conditions. The Von Mises stress and minimum safety factor considering the confining pressure effect are higher than the value without considering the confining pressure effect, while the Von Mises strain is opposite. In addition, regardless of whether the confining pressure effect is considered or not, the maximum Von Mises stress and strain show a nonlinear increasing relationship with the charge modulus, and the minimum safety factor decreases with charge modulus. However, when analyzing the structural integrity of high charge modulus, using two evaluation methods will yield completely different evaluation results. It is believed that this research can support structural integrity analysis and design of high charge modulus SRM.

Identification of Isocyanate Number on IPDI and TDI Due to Storage Period and Their Impact on Composite Solid Propellant

Isocyanate compounds serve as a curing agent in the production of composite propellants, playing a crucial role in determining their characteristics. Due to its high reactivity with moisture, it is essential to determine the isocyanate number of stored samples after a specific duration. The investigation identified the isocyanate number of IPDI and TDI that had been held for 84 and 90 months. The FTIR analysis identified the presence of the NCO group at 2240 cm-1 for TDI and 2243 cm-1 for IPDI samples held for 84 and 90 months. The isocyanate number identification decreased by 1-2% after being stored for six months. This reduction in isocyanate quantity undoubtedly impacts the development of propellant compositions. The reduction in isocyanate content will alter the propellant formulation, leading to propellants with different characteristics due to its impact on the curing ratio (NCO/OH). The curing ratio greatly impacts the mechanical properties of composite propellant.

Effects of Wire-Wrapping Patterns and Low Temperature on Combustion of Propellant Embedded with Metal Wire

Incorporating silver wires into propellant has emerged as a highly effective strategy for enhancing propellant burning rates, a technique extensively deployed in the construction of numerous fielded sounding rockets and tactical missiles. Our research, employing a multi-faceted approach encompassing thermogravimetric-differential scanning calorimetry measurements (TG-DSC), combustion diagnoses, burning rate tests, and meticulous collection of condensed combustion products, sought to elucidate how variations in silver wire quantity and winding configuration impact the combustion properties of propellants. Our findings underscore the remarkable efficacy of double tightly twisted silver wire in significantly boosting propellant burning rates under ambient conditions. Moreover, at lower temperatures, the reduced gap between the propellant and silver wire further magnifies the influence of silver wire on burning rates. However, it is noteworthy that the relationship between burning speed and combustion efficiency is not deterministic. While a smaller cone angle of the burning surface contributes to heightened burning rates, it concurrently exacerbates the polymerization effect of vapor phase aluminum particles, consequently diminishing propellant combustion efficiency. Conversely, propellants configured with sparsely twinned silver wires exhibit notable enhancements in combustion efficiency, despite a less pronounced impact on the burning rate attributed to the larger cone angle of the burning surface. Remarkably, these trends persist at lower temperatures. Based on the principle of heat transfer balance, a theoretical model for the combustion of propellants with wire inserts is developed. The reliability of this theoretical model is validated through a comparison of calculated values with experimental data. Our research outcomes carry significant implications for guiding the application and advancement of the silver wire method in solid propellants for solid rocket motors, offering valuable insights to inform future research and development endeavors in this domain.

Effect of polyol bonding agent on the mechanical properties of hydroxyl‐terminated block copolyether based composite propellant with low aluminum content

AbstractA polyol bonding agent (BA) was utilized to enhance the mechanical properties of propellant with low aluminum content. Through scanning electron microscope (SEM) and X‐ray photoelectron spectroscopy (XPS) analysis, it was discovered that BA can physically bond to the surfaces of ammonium perchlorate (AP) and hexogen (RDX). The results of mechanical testing showed that the use of BA can significantly improve the mechanical properties of propellants. Specifically, as the amount of BA increases, the maximum tensile strength (σm) of the propellant also increases, the fracture elongation (εb) initially increases and then decreases, the “dewetting ratio” continuously decreases, and the creep resistance of the propellant also increases. At a BA content of 0.2%, the σm of the propellant reaches 0.6 MPa, the εb exceeds 50%, and the “dewetting ratio” is less than 1.1.

Curing and Cross-Linking Processes in the Poly(3,3-bis-azidomethyl oxetane)-tetrahydrofuran/Toluene Diisocyanate/Trimethylolpropane System: A Density Functional Theory and Accelerated ReaxFF Molecular Dynamics Investigation.

The physical and chemical properties of solid propellant are influenced by the composition and structure of the binder, with its network structure being formed through curing and cross-linking reactions. Therefore, understanding the mechanisms of these reactions is crucial. In this study, we investigated the curing and cross-linking mechanisms of poly(3,3-bis-azidomethyl oxetane)-tetrahydrofuran (PBT), toluene diisocyanate (TDI), and trimethylolpropane (TMP) using a combination of density functional theory (DFT) calculations and accelerated ReaxFF molecular dynamics (MD) simulations. DFT calculations revealed that the steric effect of the -CH3 group in TDI exerts a significant influence on the curing reaction between TDI and PBT. Additionally, in the cross-linking process, the energy barrier for TDI reacting with TMP was found to be much lower than that for TDI reacting with the PBT-TDI intermediate. Subsequently, we conducted competing reaction processes of TMP/TDI-PBT-TDI cross-linking and TDI-PBT-TDI self-cross-linking using accelerated MD simulations within the fitted ReaxFF framework. The results showed that the successful frequency of TMP/TDI-PBT-TDI cross-linking was substantially higher than that of TDI-PBT-TDI self-cross-linking, consistent with the energy barrier results from DFT calculations. These findings deepen our understanding of the curing and cross-linking mechanisms of the PBT system, providing valuable insights for the optimization and design of solid propellants.

Micro-flow properties of NEPE propellant in a rotating depressurization combustion environment of solid rocket motor

A profound understanding of microflow processes within the combustion chamber is paramount for optimizing engine performance and ensuring safe operation in the realm of solid rocket propulsion. Specifically, the scrutiny of Nitrate Ester Plasticized Polyether (NEPE) propellant under rotating depressurization combustion conditions has garnered substantial scholarly interest. The non-uniform distribution of propellant constituents and the ensuing heterogeneity in combustion rates convolve fluid flow dynamics within the chamber, ultimately engendering intricate flow patterns, turbulent phenomena, and rotational effects. Hence, conducting numerical simulation studies assumes an imperative role in comprehensively analyzing and apprehending the intricate microflow characteristics under such demanding circumstances. This paper endeavors to accomplish this overarching objective by introducing a novel kinetic mechanism in the gas phase that meticulously accounts for complicated reactions between distinct oxidizing powders and fuel binders, representing an uncharted territory in the existing literature. Additionally, this study takes into meticulous consideration the influence of radiation effects and non-planar moving surfaces, as well as other pivotal components. Upon establishing the combustion model, exponential depressurization is induced within the gas phase, while the computational domain is deflected to accurately simulate the rotational effects. A series of methodical investigations is subsequently undertaken to validate the proposed numerical framework, yielding commendable predictive accuracy. Ultimately, this research meticulously examines the minute intricacies of microflow characteristics under such complex transient combustion conditions.

Characteristic Property of Combustion and Internal Ballistics of Triple-Based Propellant according to Particle Size of RDX

The important factors in the design of the gun propellant are impetus, flame temperature and pressure. In this paper, we considered a nitrocellulose based propellant composition that replaced sensitive NG(Nitroglycerin) with RDX(Cyclotrimethylenetrinitramine) and DEGDN(Diethylene glycol dinitrate) which high energy and low sensitivity. Particle size and content of RDX are the two main factors that affect the burning stability of RDX-based propellants. Among them, the characteristics of the propellant according to the particle size of RDX were confirmed. The relative combustion rate(R.Q., Relative Quickness) of the propellant changed according to the RDX particle size, and internal ballistics of properties of propellant were also varied. The particle size of RDX can be confirmed as a major factor in the combustion and internal ballistics characteristics of the propellant.

Effect of particle/matrix interface parameters on mechanical properties of high-energy propellant based on three-dimensional RVE model

Abstract For composite solid propellant, the interface is the medium to transfer the load between particles and matrix. Because there are a large number of particle/matrix interface layers in the propellant, the mechanical properties of the interface become an important factor affecting the macro-mechanical properties of the propellant. It is necessary to study the influence of different interface mechanical parameters on the mechanical properties of high-energy propellants. In this paper, the three-dimensional representative volume element (RVE) model of solid propellant is established, and the three-dimensional bilinear cohesion interface parameters are established for this model. The influence law of interface parameters on its mechanical properties is obtained through simulation research. The results show that both the interfacial strength and the maximum failure displacement of the interface affect the tensile strength of high-energy propellant, and the interfacial strength plays a leading role and is very significant.

Combustion and impact performance of DEGDN propellants under different size conditions

DEGDN propellant plays an important role in the use of large-caliber artillery, and its combustion and mechanical properties determine the power of large-caliber artillery. In order to study the combustion performance and mechanical properties of DEGDN propellant, four kinds of single-hole propellant with different pore diameters were prepared from DEGDN raw material ZY-11 pellets, with pore diameters of 3mm, 4mm, 5mm, and 6mm, which were recorded as 45/1-3mm, 45/1-4mm, 45/1-5mm, and 45/1-6mm, respectively. Analyze its combustion performance using closed bomb and test its impact resistance using a pendulum impact tester. The experimental results show that in the combustion process, the combustion pressure of the propellant in the higher temperature conditions, the faster its growth rate, while the whole combustion process is also faster; combustion process combustion speed will not change with the change of the aperture, only the different temperature conditions will make the combustion speed change, the higher the temperature of the maximum combustion speed it can achieve the larger. When the pore diameter is 4mm, the impact strength reaches a maximum of 30.6KJm-2, which is nearly 40% stronger than 45/1-3mm and 45/1-6mm.

Multiscale damage evolution and fracture for thermal aging energetic material based on mesoscopic observation and molecular dynamics simulation

As a composite energetic material composed of filler and matrix, the mechanical properties of HTPB propellants change constantly due to the aging of each component and component interfaces during long-term storage. However, the study of a single component or performance parameter limits disclosure of the complex aging mechanism and mechanical property evolution. Therefore, based on accelerated thermal aging experiments, this study integrated gas detection, contact angle testing, in situ loading experiments, digital image correlation numerical analysis, and all-atom molecular dynamic simulations to explore the underlying mechanism of structural morphology and mechanical property changes during propellant aging. The results showed that the aging reaction was dominated by degradation chain breaking in the early stage and oxidative crosslinking in the later stage. Degradation chain breaking enhanced interfacial bonding between the two phases, while oxidative crosslinking deteriorated interfacial bonding. Compared with lossless specimens, short-term thermal aging weakened the strength of the matrix and enhanced bonding interfaces. During in situ loading, interfacial pore expansion and matrix damage and fracture occurred simultaneously and the matrix became the key damage source. However, with the extension of thermal aging time, the strength of the matrix increased, bonding interface weakened, and bonding interfaces became a weak point of loading. This made the high-strain area change from the matrix at the initial loading stage to the bonding interface and the damage of the bonding interface caused whole fracture of the specimen. The external loading resulted in the rearrangement of interfacial molecular chains, change of molecular spacing, and initiation and expansion of micropores between molecular chains. The tractor-separation curve conformed to the exponential cohesion model and the damage mode of the interface layer changed between adhesion and cohesion due to thermal aging.

Synchronized measurement method of burning rate and combustion temperature of a solid propellant specimen

To characterize the combustion properties of solid propellants, the synchronized measurement method of burning rate and combustion temperature is proposed combined shadow imaging and radiation imaging. Using spectroscopic and filtering imaging, shadow and radiation images of a solid propellant specimen are obtained synchronously. Burning rate is calculated by burning surface movement velocity of shadow images, and combustion temperature is calculated by radiation image thermometry. Measurement accuracies of burning rate and combustion temperature of the solid propellant specimen are validated by other independent measurement methods. On this basis, the synchronized measurements of burning rate and combustion temperature of different formulations of solid propellant specimens under different working conditions are carried out. The results show that the influence on burning rate and combustion temperature of pressure and formulas is different. Therefore, the synchronized measurement of burning rate and combustion temperature can provide more direct data support for the evaluation of solid propellant combustion performance.

Microencapsulation of ADN with HTPB‐based membrane in the presence of the bonding agents Tepan or Tepanol**

AbstractAmmonium dinitramide (ADN) has appeared as a promising oxidizer for green propellants and thereby a potential substitute for ammonium perchlorate, largely in use in composite propellants for tactical and strategic long‐range missiles. The novelty lies in replacing ammonium perchlorate with a chlorine‐free oxidizer less harmful to the health and environment. However, ADN is hygroscopic and can potentially react with other chemical components, which could be overcome by microencapsulating the particles. The simple coacervation method was tested herein to microencapsulate ADN with a membrane made of hydroxyl‐terminated polybutadiene as pre‐polymer and methylene diphenyl diisocyanate as the curing agent. The effect of polyamine bonding agents on the capsule formation was tested by adding 0.5 or 2 % of Tepan or Tepanol, whose efficacy to bond to ADN was confirmed by detecting ammonia release through infrared spectroscopy. The capsule membrane was examined by optical and scanning electron microscopy. The dissolution time and rate were the parameters adopted to quantify permeability in a straight dissolution test in water, which demonstrated that 0.5 % Tepanol can provide the most effective protection. The infrared spectroscopy indicated that 60 °C temperature for prolonged periods, normally experienced by propellants, does not chemically affect the capsules’ membrane but can turn it lumpy. In conclusion, these polyamine bonding agents can assist the capsule formation over ADN particles using the simple coacervation method, however, their functionality on mechanical properties of propellants needs to be substantiated in forthcoming works as well as the effect of the concentration of bonding agents on propellant formulations.

Impacts of metal oxide crystalline structure on the decomposition of solid propellants under combustion heating rates

Ammonium perchlorate (AP) has been the oxidizer of choice for composite solid propellants for decades and has been the object of decomposition studies for safety monitoring. Typically, studies perform differential scanning calorimetry (DSC) or thermogravimetric analysis (TGA) to show how metal oxides (MO) commonly incorporated into propellants alter AP decomposition rates and completeness of reaction. Most past decomposition work studies temperatures below the crystal phase transition (240°C) from orthorhombic phase to cubic, which impose internal stresses within the lattice of AP particles. Phase change induces a partial decomposition, which does not follow a shrinking core behavior; instead, it develops a network of pores up to a few microns in size throughout. During low-temperature decomposition, particles lose approximately 30-40 % of their mass. Simultaneous DSC/TGA (STA) use low heating rates for combustion environments but offer information on MO's catalytic and electron transport modifications. This work demonstrates how the crystalline morphology MO additives mechanistically alter the combustion properties of composite propellants by identifying modes for promoting decomposition in AP and HTPB. MO morphology determines the electron structure of the molecule, which sets the band gap properties of the particles. This study evaluated different morphologies and a range of MOs to identify the most active MO for decomposing A.P. Kissinger analysis was applied using STA data at heating rates of 10-, 20-, and 30-°C/min revealing significant shift in the high-temperature decomposition with a range of metal oxides. Higher heating rates were evaluated using CO2 laser ignition to identify the time to first gas using the MO that offers the best heat absorption, such as the aluminum oxides. This higher heating rate was found do more accurately represent the changes in the combustion rates in the final propellant mixture. Additionally, it was shown that electron transport additives like CuO show the most significant impact on combustion, and thermal absorbers offer the lowest impact on AP/HTPB combustion. This understanding offers a new approach to propellant design not previously presented and suggests future testing for tailoring the use of metal oxides in propellants.

Molecular dynamics simulation of bonding role of cyclic borate ester in improving mechanical properties of the HTPB propellant

Excellent mechanical properties are the fundamental requirement for the practical application of solid propellants. However, solid oxidant fillers such as hexahydro-1,3,5-trinitros triazine (RDX) seriously weaken the mechanical strength of the propellant due to poor adhesion with the binder matrix. In this work, novel cyclic borate ester (CBE) bonding agents with long-chain alkyl moiety were designed by the molecular dynamics (MD) method to improve the unfavorable interfacial bonding effect. The interface models, which consist of RDX substrate and binder matrix where hydroxyl-terminated polybutadiene (HTPB) crosslinked with curing agents, were constructed to reveal the interface enhancement mechanism by CBE bonding agents. The results indicate that the designed CBE bonding agents with appropriate solubility parameters (SPs) can migrate from the binder matrix to the RDX surface. Besides, CBE bonding agents exhibit strong interaction with RDX compared to the binder matrix due to the formation of hydrogen bonds. Finally, the simulated tensile tests with two tensile modes showed that CBE bonding agents can enhance the interfacial strength and reduce stress softening rates by the calculation of debonding work. Moreover, this improvement can be controlled by adjusting the chain length of CBE bonding agents. The results obtained in this work deepen the understanding of the interfacial strengthening mechanism of bonding agents for solid propellants and provide a framework for designing high-performance bonding agents based on the MD method.

Novel graphene iron organic nanocomposites for enhancing combustion and safety properties of AP-HTPB propellant

Graphene iron tannate and graphene iron gallate nanocomposites (G-D-Fe and G-M-Fe) were designed, synthesized and characterized to deal with the performance requirement of AP-HTPB propellent. Effects of the G-D-Fe and G-M-Fe on the thermal decomposition of AP, laser ignition, combustion and safety characteristics of AP-HTPB propellant were explored. By taking Catocene as the benchmark (156.80 mJ), the G-D-Fe improves the safety performance of propellant, and the ignition energy (E50) is increased by 17.25 mJ. Besides, the G-D-Fe and G-M-Fe have positive effects on enhancing the burning rate and the reducing the pressure exponent of propellant, and the influence of G-D-Fe is more significant. The burning rate enhances from 13.87 to 16.92 mm·s−1 at 15 MPa and the pressure exponent reduces from 0.47 to 0.32 in 4∼15 MPa after the introduction of G-D-Fe. The catalytic mechanism of G-D-Fe were analyzed on the basis of flame morphology, combustion wave structure, in-situ FTIR and gas-phase FTIR-MS. The outstanding catalytic property of G-D-Fe is due to its promotion effect on AP pyrolysis and the formation of stable products, which is conductive to the intense reactions of gas products and the thermal transmission between burning surface and gas phase, and reflecting in the significant enhanced burning rate.