Burning Surface Of Propellant Research Articles

In situ preparation of aluminum-copper pentafluorobenzoate energy microspheres: Enhancing aluminum combustion properties in composite propellants

Aluminum powder is the main metal fuel in composite propellants, and upgrading the proportion of aluminum powder loaded in solid propellants is currently one of the most effective means of increasing propellant energy. Nonetheless, the substantial disparity in boiling points between the core aluminum particles and the alumina layer enveloping their surfaces often results in reduced reactivity, inadequate combustion, and combustion agglomeration during the combustion process of the aluminum powder. Herein, copper pentafluorobenzoate (CuPF) interfacial layer was established on the surface of aluminum powder through the self-assembly coordination reaction between pentafluorobenzoic acid and copper ions, and Al-CuPF energy microspheres were successfully prepared. A comprehensive examination was carried out to scrutinize the surface morphology, elemental composition, and distribution of the Al-CuPF energy microspheres. Furthermore, meticulous studies were conducted on the ignition and combustion properties of these Al-CuPF energy microspheres under various atmospheric conditions. It was found that the Al-CuPF interfacial layer within the Al-CuPF energy microspheres plays a pivotal role in enhancing ignition initiation and acting as a catalyst to facilitate the more efficient combustion of aluminum powder. Moreover, the combustion behavior of Al-CuPF energy microspheres when integrated into composite propellants has been exhaustively researched. Notably, composite propellants incorporating Al-CuPF energy microspheres exhibit heightened burning temperatures and an increased burning rate. Additionally, the combustion and agglomeration dynamics of Al-CuPF energy microspheres on the propellant burning surface were meticulously documented using a high-speed camera. The findings revealed that Al-CuPF energy microspheres spend a significantly shorter duration on the propellant burning surface and notably mitigate the occurrence of aluminum agglomeration issues thereon. This research contributes fresh insights into the application of innovative aluminum-based fuels within the realm of composite propellants.

Influence of nickel powders on burning behaviors of single-based propellant in variable-pressure and constant-pressure combustion conditions

Single-based propellants containing 0.4 wt%, 0.8 wt% and 1.6 wt% nickel powders(NPs) were prepared and investigated in closed bomb tester and transparent combustion chamber, which can provide variable-pressure and constant-pressure conditions for propellant burning, respectively. In closed bomb tester, propellants burned to a pressure above 5–220 MPa. The results showed that burning rate of propellants was increased by 19.0 %, 12.5 % and 11.9 % in 5–8 MPa, 8–16 MPa and 20–200 MPa, when 1.6 wt% NPs used in propellant. Meanwhile, propellants burning in combustion chamber were with a constant pressure of 1.0 MPa, 1.5 MPa, 2.0 MPa, 3.0 MPa and 5.0 MPa. The burning rate of propellants can be increased by NPs in 1.0–5.0 MPa. While the catalytic efficiency(Z) had high dependence on NPs contents and burning pressure. At 5.0 MPa, 0.4 wt%, 0.8 wt% and 1.6 wt% of NPs enhanced burning rate of propellants by 42 %, 44 % and 68 % than that of original propellants. In pressure below 3.0 MPa, burning rate of propellants cannot be increased with 0.4 wt% and 0.8 wt% NPs, but that can be enhanced with 1.6 wt% NPs, suggested that higher NPs content and higher pressure were conducive to catalytic effect, which was significant with the pressure above 5 MPa. Furthermore, after NPs incorporating into propellant, the distance from flame zone to propellant burning surface was significantly reduced. And the potential catalytic mechanism was proposed.

Modeling of aluminum agglomeration at solid propellant burning surface by combining aluminum ignition with LDP-SC cluster analysis

Aluminum particles at burning surface may agglomerate during the combustion process of composite solid propellant, involving the aggregation and ignition of aluminum particles as well as the secondary coalescence of agglomerates. So far it is still a huge challenge to capture the agglomeration process and obtain the size distribution of agglomerates accurately. Modeling of aluminum agglomeration was carried out by constructing several submodels, including propellant combustion, cluster analysis of aluminum particles, aluminum ignition and coalescence based on the packed three-dimensional propellant microstructure unit. Dynamic combustion process of propellant microstructure unit was simulated firstly, which characterizes the temperature distribution above the burning surface. Next, aluminum clusters were grouped using a new local density peaks-spectral clustering (LDP-SC) algorithm for initial aluminum particles in microstructure unit. The ignition of aluminum particles in clusters exposed from burning surface was then examined by considering their distributions in the flame of propellant. These exposed aluminum particles in a cluster yield an agglomerate once an aluminum particle is successfully ignited. Secondary mergence of multiple agglomerates was also considered when the distance between two particles in different agglomerates was less than a critical separation distance. With these submodels, the agglomeration process of a typical propellant and properties of aluminum clusters and agglomerates were estimated. The predicted size distribution of agglomerates shows reasonable agreement with experimental data, indicating that the present model is capable of predicting aluminum particle agglomeration during composite propellant combustion.

Experimental and numerical investigation of aluminum particle combustion driven instability in solid rocket motors

Aluminum particles are extensively employed to enhance the energetic properties of propellants. However, the heat release rate (HRR) fluctuations resulting from aluminum combustion can potentially affect the stability of solid rocket motors (SRMs). Using a self-made high-pressure combustion diagnostics system and laser ignition system, this work investigates the combustion behavior of single aluminum particle under propellant atmosphere and acoustic excitation. An unsteady aluminum combustion model using Ranz-Marshall correlation and involving the heat and mass transfer equation was developed. The D2 law of aluminum particle combustion in a real propellant under 0.5 MPa was fitted as t = D2/0.31. The measured combustion rate of laser-ignited aluminum particles was found to increase from 0.381mm2/s to 0.465 mm2/s when an acoustic excitation of sine wave(200 Hz, 20Pa) was applied to the propellant, which agreed well with the prediction of the proposed unsteady combustion model whose error was 4.7%. Then, the verified correlation was included in the SRM to calculate the HRR fluctuation of the aluminum on acoustic boundary near the propellant burning surface. The effects of particle diameter, gas velocity, oscillation amplitude and frequency on the HRR fluctuation were obtained. It was found that the HRR fluctuation increases with the increasing particle diameter, gas velocity, and oscillation amplitude, and that due to the acoustic boundary effect, the HRR fluctuation is more sensitive under low-frequency oscillation. Based on the HRR fluctuation of aluminum combustion and particle damping, the growth rate of instability resulting from aluminum combustion in an SRM was investigated. The results showed that the growth rate of instability resulting from aluminum combustion were 9.38s-1 and 24.06s-1, while the particle damping (−25.8s-1) caused by combustion residuals, which suggests the aluminum particles play a critical role in SRMs combustion instability.

The characteristics and mechanism in combustion extinguished process based on self-extinguished solid propellants

With the addition of the quaternary ammonium salt used as surfactant, a self-extinguished solid propellant was prepared. The combustion performance, extinguished burning surface and thermal analysis of the self-extinguished propellant were studied. The quaternary ammonium is a key point to the extinguished process in combustion because of the influence in interface between oxidizers and binders. The similar critical pressure of extinguished and the inflection point of the pressure exponent indicated that the mechanism in combustion extinguished process was also related to the reaction process of oxidizers.

Agglomerate Size Evolution in Solid Propellant Combustion under High Pressure

Solid propellant combustion and flow are significantly affected by condensed combustion products (CCPs) in solid rocket motors. A new aluminum agglomeration model is established using the discrete element method, considering the burning rate and formulation of the propellant. Combining the aluminum combustion and alumina deposition model, an analytical model of the evolution of CCPs is proposed, capable of predicting the particle-size distribution of completely burned CCPs. The CCPs near and away from the propellant burning surface are collected by a special quench vessel under 6~10 MPa, to verify the applicability of the CCP evolution model. Experimental results show that the predicted error of the proposed CCP evolution model is less than 8.5%. Results are expected to help develop better analytical tools for the combustion of solid propellants and solid rocket motors.

Conjugate heat and mass transfer at heterogeneous burning surface of AP/HTPB/Al composite propellant at 2–10 MPa

Conjugate heat and mass transfer at heterogeneous burning surface has significant effect on combustion characteristics of AP/HTPB/Al composite propellant. In the present study, the conjugate heat and mass transfer characteristics of AP/HTPB/Al composite propellant are studied experimentally and numerically based on formulated combustion model and experimental facilities. With the increase of pressure and Al content, flame temperature increases significantly and leads to stronger heat transfer to the burning surface. With the content of Al varying from 18 wt% to 9 wt%, the total heat transfer coefficient decreases for all the sampled positions of the burning surface with a maximum decrease of 12% at the AP/HTPB interface. It is found that heat convection accounts for more than 97% of the total heat transfer coefficient. However, with pressure increasing from 2 MPa to 10 MPa, the total heat transfer coefficient and the respective contribution of heat convection and radiation vary slightly. Owing to the heterogeneous sandwich-like structure of the propellant, the temperature distribution at the burning surface of propellant is highly non-uniform. Therefore, heat transfer characteristics vary significantly at different positions of the burning surface. The heterogeneity of the burning surface is exacerbated by the non-uniform distribution of heat transfer coefficients at different positions of the burning surface. The burning rate of the propellant varies from 0.5 mm/s to 9.5 mm/s for different positions of the burning surface at 10 MPa. The results of micro-scale flame structure show that the mass transfer of fine Al particles in the combustion of AP/HTPB/Al composite propellant can be basically divided into four stages, i.e., migration, melting, agglomeration and oxidation/combustion. Coral reef-like aggregate structure is formed by some of the fine Al particles before oxidation and combustion. Increasing pressure is conducive to avoiding formation of large Al agglomerates and enhancing mass transfer at the heterogeneous burning surface.

Numerical simulation of aluminum particle agglomeration near the burning surface of solid propellants

Aluminum is an important additive in solid propellants used to improve energy efficiency, but agglomeration of aluminum particles can reduce the specific impulse of the engine, increase two-phase flow losses and cause ablation of the adiabatic layer. Therefore, it is important to understand the mechanism of formation, mode of movement and particle size characteristics of aluminum agglomerates in solid propellants to study the inhibition of agglomeration. In this paper, a three-dimensional numerical simulation of aluminum particle agglomeration near the burning surface of solid propellants was conducted using the discrete unit method based on direct tracking calculations of the detailed motion of individual particles. The model considered physical processes such as burning surface motion, aluminum particle precipitation, turbulent pulsation and the heating and melting of aluminum particle and used a grid-based contact detection method to accelerate the computational process. The agglomeration process of a solid propellant with specific formulation was experimentally investigated. Aluminum particle aggregates underwent splitting, and the final aluminum particle agglomerates were spherical droplets with typical structure. The agglomeration process of this propellant was investigated by means of a simulation. The accuracy of the agglomeration model was verified by comparing the results of the model for the agglomeration process and the shape and size distribution of agglomerates with available experimental data. Compared with the experimental results, the deviations of the equivalent particle sizes of agglomerates D10, D50, D90, D4,3, and D3,2 obtained by using the simulation results were 6.3%, 6.3%, 9.4%, 7.5% and 6.7%, respectively. The aluminum particle agglomeration model established in this paper can predict the particle size distribution of aluminum agglomerates near the burning surface of solid propellants. The model was also used to further investigate the effects of aluminum particle content, ammonium perchlorate (AP) particle size and environmental pressure on aluminum particle agglomeration. Increasing the aluminum content or AP particle size aggravated aluminum particle agglomeration, whereas increasing pressure weakened it.

Numerical analysis of 3-D inner flow field for ladder-shaped multiple propellant rocket motor

In order to study the characteristics of ladder-shaped multiple propellant rocket motor, it was conducted that the numerical simulation of the full 3-D inner flow field for ladder-shaped multiple propellant rocket motor working process using FLUENT calculation software. The secondary development was carried out through UDF interface programming, and the changing rules of the added mass in propellant burning surface and dynamic boundary with the burning rate were set, so that the dynamic process of propellant combustion is accurately simulated. The numerical simulation results demonstrated the spatial distribution information of parameters, such as pressure and flow rate at each moment in the inner flow field, and the pressure-time curves of three monitoring points in front, center and rear of the combustion chamber. Then, several key characteristic moments were selected for specific analysis, and the pressure-time curve of the monitoring point was compared with the experimental data to show that the calculated results were basically accurate. The pressure, flow rate and the pressure difference between the internal and external surfaces of the propellant in the rocket motor inner flow field were analyzed. At present, the dynamic boundary treatment method that the burning rate of the propellant varies regularly with the surface pressure has not been applied in the 3-D calculation model. The research results of this paper can provide reference for the numerical simulation of similar structure of solid rocket motors.

On the Use of Fluorine‐Containing Nano‐Aluminum Composite Particles to Tailor Composite Solid Rocket Propellants

AbstractThe burning rate of solid propellants is an important factor for optimizing rocket motors and improving performance. The enhanced burning rate can increase thrust and reduce a propulsion system‘s overall size and weight. In this study, a novel nano‐aluminum/THV composite additive was prepared and introduced into a solid ammonium perchlorate/polybutadiene composite solid rocket propellant to enhance its burning rate. The morphology of the composite particle additive and its effects on combustion were characterized. The use of small quantities (<15 wt.%) of the additive resulted in a burning rate enhancement of up to 2.1 times that of the conventional coarse aluminized propellant with a specific impulse loss of only 3 seconds, and as much as 4.7 times enhancement with a predicted loss of 22 seconds in theoretical specific impulse. Some of this loss may be recovered by the improved combustion efficiency in smaller rocket motors because the additive was shown to significantly reduce the aluminum agglomeration at the propellant burning surface and reduce the size of reaction products which may reduce two‐phase flow losses. The additive also provides wide burning rate tailorability, favorable for motor, grain, and thrust curve design. The burning rate enhancement mechanism is thought to be a physical cratering mechanism governed by the burning rate disparity between the binder/oxidizer system and the nano‐aluminum/fluoropolymer additive and not a chemical catalytic effect.

Burn rate of a novel boron-AN-water green solid propellant

A novel composition of green solid rocket propellant has been suggested and examined. Boron powder was chosen to serve as fuel, for its very high energetic potential, which exceeds that of other common metals, as well as that of fossil fuels. As an oxidizer, the energetic carbon-free and chlorine-free ammonium nitrate (AN, NH4NO3) was considered, and distilled water was used as binder. The mass-based composition consisted of ∼20% boron, ∼60% AN and ∼20% water (“20/60/20″), and small amount of gelling agent was added to enhance the propellant solidity and mechanical properties. Atmospheric burn rate experiments resulted in unstable combustion and flame quenching, probably due to poor heat feedback from the flame to the propellant burning surface. Pressurized experiments at a range of 0.8–4.3 MPa (using inert nitrogen) exhibited steady combustion and obeyed well to the power-law behavior of burning rate with pressure, following the equation: r˙[mm/s]=(2.41±0.22)·(P[MPa])0.84±0.11. Although the measured regression rates were similar to those seen in conventional composite solid propellants for the investigated pressure range, the obtained pressure exponent (n = 0.84) is considered high for solid rocket motor use. Eventually, combustion products analysis revealed that unlike many past cases where boron was not fully combusted in solid propellants, metal oxidation was complete and boron oxide (B2O3) was practically the sole reaction product of boron.

Combustion and agglomeration characteristics of boron particles in boron-containing fuel-rich propellant

The energy release efficiency of boron (B)-containing fuel-rich solid propellants depends largely on the combustion behavior of the B particles contained therein. Here, the combustion process of B particles was first studied using a laser ignition testing system. Then, the combustion and agglomeration characteristics of B particles in a fuel-rich propellant were further investigated. To this end, a microtube burner was specially manufactured to mimic the secondary combustion of boron-containing propellant. Results show that the primary combustion intensity of the boron-containing propellant was relatively weak, and the combustion temperature was 1720±32 K. B particles would participate in the primary combustion reaction of the propellant, resulting in short time detection of the characteristic green flame and BO2 spectrum. Severe self-conglomeration of B particles occurred on the propellant burning surface, forming irregular coral-like B particle conglomerates. The development of the quasi-secondary (the word “quasi” is added to highlight the difference between the experimental conditions and the actual conditions in engines) combustion flame showed that hot B particles were an important ignition source for the gaseous fuel and that the existence of excess gaseous fuel in the primary plume was an important prerequisite for self-sustaining combustion of B particles in air. The emission spectra of propellant quasi-secondary combustion were very similar to those of primary combustion, but the intensity was significantly increased. Microstructures of the combustion residues showed that the oxide layer on the surface of B particles played an important role during the formation of B conglomerates. Irregular B conglomerates detached from the propellant burning surface might further develop into spherical B agglomerates in the high-temperature secondary combustion flame zone.

New Design Method of Solid Propellant Grain Using Machine Learning

The correlation between solid propellant grain configuration and burning surface area profile is a complicated nonlinear problem. Nonlinear optimization has been adopted to design grain configurations that satisfied the objective area profiles. However, as conventional design methods are impractical, with limited performance, it is necessary to investigate alternatives. Useful information for grain design can be obtained by analyzing the aforementioned correlation. However, this aspect has not been studied owing to the requirement of large amounts of data and analysis techniques. In this study, machine learning was used to develop a new design method. The objective of machine learning was to train a model to classify classes of data. The database stores various sets of configuration variables and their classes. The proposed Gaussian kernel-based support vector machine model predicts the class of newly designed grains. The results verified that the model accurately predicted the class of the set of configuration variables and can be used to modify the set of configuration variables to satisfy the requirement. Thus, it was confirmed that machine learning is an appropriate approach to grain design; however, further research is needed to analyze its practicality.

Studies on Aluminum Agglomeration and Combustion in Catalyzed Composite Propellants

Composite propellants are tested using the quench particle collection bomb (QPCB) for the pressure ranging from 2 to 8 MPa to estimate the particle size distribution of aluminum agglomerates from quenched combustion residues emerged out from the burning surface. The major ingredients included in the propellants are ammonium perchlorate (AP), aluminum (Al), hydroxyl-terminated polybutadiene (HTPB), and toluene diisocyanate (TDI). Five propellant compositions are considered in this study; two of them are mixed with catalysts. Propellant formulation variables like the coarse AP/fine AP ratio, total solid loading, catalyst percentage, and aluminum content are varied to assess their effects on the aluminum agglomeration process at different pressures. Unburnt aluminum in agglomerates is continuously getting combusted as they move out from the propellant burning surface. Large agglomerates comprise both Al2O3 and unburnt aluminum. The majority of agglomerates are spherical in shape, and the sizes vary from 31 to 115 $$\mu$$ m for non-catalyzed propellants and from 28 to 136 $$\mu$$ m for catalyzed propellants over the tested pressure conditions. These results can give further insight into the aluminum agglomeration process of catalyzed and non-catalyzed propellants and also affect the choice of the propellant ingredient percentage aimed at reducing aluminum agglomeration, which causes two-phase flow losses of thrust and slag accumulation in full-scale solid rocket motors.

Characterization of the influence of aluminum particle size on the temperature of composite-propellant flames using CO absorption and AlO emission spectroscopy

Understanding the temperature of aluminized, composite-propellant flames is critical to achieving robust rocket motor designs and developing accurate, predictive models for propellant combustion. This work presents measurements of (1) the temperature of CO (within the flame bath gas) and (2) the temperature of AlO (located primarily within regions surrounding the burning aluminum particles) within aluminized, composite-propellant flames as a function of height above the burning propellant surface. Three aluminized, ammonium-perchlorate (AP), hydroxyl-terminated polybutadiene (HTPB) composite propellants with varying aluminum particle size (nominally 31 μm, 4.5 μm, or 80 nm) and one non-aluminized AP-HTPB propellant were studied while burning in air at 1 atm. A wavelength-modulation-spectroscopy technique was utilized to measure CO temperature and mole fraction via mid-infrared wavelengths and a conventional AlO emission-spectroscopy technique was utilized to measure the temperature of AlO. The bath-gas temperature varied significantly between propellants, particularly within 2 cm of the burning surface. The propellant with the smallest particles (nano-scale aluminum) had the highest average temperatures and far less variation with measurement location. At all measurement locations, the average bath-gas temperature increased as the initial particle size of aluminum in the propellant decreased, likely due to increased aluminum combustion. The results support arguments that larger aluminum particles can act as a heat sink near the propellant surface and require more time and space to ignite and burn completely. On a time-averaged basis, the temperatures measured from AlO and CO agreed within uncertainty at near 2650 K in the nano-aluminum propellant flame, however, AlO temperatures often exceeded CO temperatures by ≈ 250 to 800 K in the micron-aluminum propellant flames. This result suggests that in the flames studied here, and on a time-averaged basis, the micron-aluminum particles burn in the diffusion-controlled combustion regime, whereas the nano-aluminum particles burn within or very close to the kinetically controlled combustion regime.

Thermo-physical Properties and Combustion Wave Aspects of RDX Contain Low Aluminium Composite Propellant

Thermo-physical properties play an important role in combustion and heat transfer of propulsion system. The thermo-physical properties (thermal diffusivity, specific heat and thermal conductivity) aspects were studied on incorporation of RDX by partial replacement of AP in low aluminized propellant. These properties were experimentally determined by Flash technique at ambient temperature up to 100 °C. The finding reveals that on increasing the ratio of RDX to AP, there is an increase in specific heat and thermal conductivity while thermal diffusivity decreases, accordingly. The effect of RDX on combustion mechanism is evident in combustion wave, which was clearly visible in high-speed images of propellant burning surface at ambient conditions.

Mean Burning Rate Variation in Composite Propellant Combustion Due to Longitudinal Acoustic Oscillations

This paper focuses on investigating “combustion instability,” a phenomenon that mainly involves the interaction of the propellant flames with the acoustic oscillations prevalent in full-scale rocket motors. The effect of excited acoustic pressure oscillations on the mean burning rate of solid propellants was estimated over the pressure ranges from 1 to 7 MPa using a window bomb facility coupled with a rotary valve. Both non-aluminized and aluminized composite propellants were used. A rotary valve was used to drive the acoustic oscillations at frequencies of 140, 180 and 240 Hz with the pressure amplitude perturbations ranging from 0.1 to 1.4% of mean chamber pressure. Frequency shift due to propellant combustion was also investigated for both the types of propellants. The acoustic oscillations enhance the heat transfer between the flame and propellant burning surface, altering the mean burning rate. The acoustic pressure oscillations influence the dynamics of aluminum particles and its agglomeration processes, which modifies the mean burning rate. The variations in excited frequencies show a significant impact on the mean burning rate. The burning rate augmentation factor shows that acoustic excitation is more predominant at low pressures and high frequencies, whereas it is relatively marginal at high pressures. The evaluated maximum augmentation factors are 1.45 and 1.51 for nonaluminized and aluminized propellants when compared with those without oscillations.

Ultra‐Fine Aluminium Characterization and its Agglomeration Features in Solid Propellant Combustion for Various Quenched Distance and Pressure

AbstractAddition of micrometer‐sized aluminium powder in solid propellant increases the specific impulse whereas it also contributes to thrust loss due to two‐phase flow. Aluminium particles form large sized agglomerates during propellant burning. In this context, agglomeration feature of ultra‐fine aluminium (UFAL) in solid propellant combustion is investigated using experimental study. The synthesized UFAL powder is used with the intensity weighted harmonic mean size of 438 nm which is produced by RF induction plasma method. The morphology and purity of the UFAL is verified using SEM and EDAX studies respectively. The ultra‐fine particles are also characterized using gas volumetric method, surface area measurement and XRD analysis. Quench particle collection technique is adopted to estimate the aluminium agglomerate measurement over the pressure ranges from 2 MPa to 8 MPa. The flame quenching distance from the propellant burning surface is varied from 5 mm to 71 mm to estimate the effect on agglomerate size. The inherited residual forces of UFAL strongly influences the agglomeration process, and also the aluminium agglomerate sizes vary from propellant burning surface towards the complete aluminium combustion. The XRD and XPS results of the agglomerates show that the aluminium content decreases and alumina content increases far away from the burning surface, invariably represents the initiation of combustion process over the considered quenching distances. UFAL powder exhibits significant agglomeration with the size ranging from 11–21 μm. This will substantially benefit from exhaust signature point of view and also reduction in two‐phase flow losses to thrust in contrast to micron sized‐aluminium particles in propellant which displays large sized aluminium agglomerates.

Effect of RDX content on the agglomeration, combustion and condensed combustion products of an aluminized HTPB propellant

Cyclotrimethylene trinitramine (RDX) is widely used in aluminized solid propellants to increase specific impulse. Experimental investigations on the effect of RDX content on the propellant combustion and agglomeration were thoroughly conducted based on Thermogravimetry-differential scanning calorimetry, laser ignition, combustion diagnosis and a new home-made condensed combustion products (CCPs) collection device. The results show RDX inclusion inhibits the decomposition of AP. The ignition delay and self-sustaining combustion time of aluminized propellant increase, while the combustion intensity decreases with increasing RDX content. RDX addition is found to decrease the burning rate within 6–10 MPa. Compared with the baseline propellant, RDX leads to obvious aggravation of aluminum agglomeration on the propellant burning surface. The 6 wt% and 12 wt% RDX-included propellants increase the average CCPs size from 46.3 μm to 86.7 μm and 96.6 μm, respectively. The combustion efficiency of aluminum in propellant is reduced by 15% when the RDX content is increased from 0 to 12%. Overall, RDX plays a comprehensive effect on the ignition, combustion and agglomeration characteristics of aluminized propellant. Results of this study can help guide the application and development of RDX in aluminized propellants for solid rocket motors.

Study on Ignition Delay Time of Al/Mg Fuel-rich Propellant under Tangential Air Flow

Solid propellant, as the power source of propulsion system, has obvious influence on the internal ballistic characteristics of ramjet. The ignition and combustion charac-teristics of solid propellant help to reveal the combustion mechanism and have im-portant guiding significance for estimating the performance of propulsion system. In this paper, the ignition and combustion characteristics of Al/Mg fuel-rich propellant under tangential air flow were studied. Laser ignition experiment was carried out on the Al/Mg fuel-rich propellant at different tangential air flow rates and tangential air flow components. The results show that the air flow rate and air flow composition can affect the flame shape and the brightness of the propellant. At the same time, with the increase of air flow rate, the more obvious the erosion combustion effect on the pro-pellant burning surface. The air flow rate also affects the ignition delay time. When the air flow rate is low, the ignition delay time decreases with the increase of air flow rate. When air flow rate is 0.8L/min, the ignition delay time is reduced to the mini-mum of 350ms, but with the air flow rate continues to increase, the ignition delay time is increasing. As the oxygen content in the air flow increases, the ignition delay time increases. However, when the oxygen content is less than 14% or higher than 21%, the effect of oxygen content on the ignition delay time is weakened.