In this experimental study, we explore how the combustion and agglomeration characteristics of solid propellants can be modified by replacing aluminum with ternary Al/CuO/PVDF(Polyvinylidene ...Fluoride) metastable intermolecular composites (MICs). Using thermogravimetric−differential scanning calorimetry, a laser ignition setup, and a high-pressure propellant combustion system, we examine the thermal reactivity, ignition behavior, burning rate, agglomeration, and condensed combustion products of various MICs-inclusion propellants. We find that the use of ternary Al/CuO/PVDF MICs can decrease the onset temperature and increase the oxidation efficiency of Al, as compared with common binary MICs. The ratio of CuO/PVDF is found to have a strong effect on the combustion intensity of the metallic powders, but only a weak effect on their ignition delay. Higher proportion of PVDF is found to have a negative effect on the combustion rate. Specifically, the combustion intensity is found to decrease first and then increase as the PVDF content increases. The exothermic Al/CuO reaction is found to alter the heat transfer of the condensed layer on the propellant surface, while PVDF plays a critical role in the thermal feedback of the gaseous reaction. When combined, these two factors lead to a ± 30% variation in the propellant burning rate. Furthermore, ternary MICs are found to reduce agglomeration during propellant combustion. The two optimal formulations are 2.5 wt.% CuO/2.5 wt.% PVDF and 3.5 wt.% CuO/1.5 wt.% PVDF. The use of these formulations is found to decrease the mean particle size of the condensed combustion products to only 11 μm, as compared with 76 μm for the original Al-inclusion propellant. High-speed microscope images reveal the existence of crack phenomena in the pockets surrounding the AP particles on the burning surface due to the Al/CuO reaction, as well as the floccule breakup mechanism of Al aggregates in the presence of PVDF, both of which reduce agglomeration. Overall, this study shows that replacing Al with ternary Al/CuO/PVDF MICs has significant effects on the agglomeration, combustion and ignition features of aluminized propellants. The findings of this study can be used to aid the development of highly adjustable functional catalysts for solid propellants.
The secondary combustion between unburnt fuel in rocket plume and ambient air significantly affects the stealth characteristics of rocket motors. In order to reveal the effect of secondary boundary ...layer combustion of hydrogen on rocket plume heat release characteristics, a two-dimensional axisymmetric model for boundary layer combustion between the rocket plume and air is established. The results identifies significant exothermic effect due to secondary combustion reaction between the plume and air. Compared to the nozzle plume without considering secondary combustion, the average temperature of the plume with secondary combustion increases by 47%. As hydrogen content in the nozzle plume increases from 2.7% to 4.2%, the average temperature of the nozzle plume increases by 17.2% due to enhanced combustion heat release. The secondary combustion process is weakened with the increase of flight altitude. The average temperature of the nozzle plume decreases by 11.8% from 2 km to 8 km of flight altitude. This is because both the pressure and temperature of the ambient air decrease when the flying altitude rises, which leads to lower secondary combustion reaction and heat release rate. Moreover, as the ambient pressure decreases, the hydrogen concentration in the plume decreases due to plume expansion effect. The results indicate that hydrogen content and flight altitude affect secondary combustion by different mechanisms. The hydrogen content directly affects the concentration term of the reaction rate of secondary combustion. However, the flight altitude affects the kinetic rate of secondary combustion by changing both the temperature and partial pressure of oxygen. The present study provides better insight into the interaction mechanism between nozzle plume and ambient air.
•Established numerical model for secondary combustion of rocket plume.•Higher hydrogen content enhances secondary combustion process.•Secondary combustion heat release is weakened as flight altitude increases.•Secondary combustion is controlled by shear mixing and heat transfer.
•New fluoropolymer containing hydroxyl groups is developed with high coating stability.•Ignition characteristics of single Al-FCOS composite particles is carefully examined.•The new coatings greatly ...shortens the ignition delay of virgin Al particles.•Al-FCOS enhance the combustion intensity and efficiency of common solid propellant.•Combustion and agglomeration can be tuned by adjusting the new fluoropolymer content.
Fluorine-containing organic substances (FCOS) are often added to aluminized solid propellants to improve their ignition and combustion characteristics, but existing FCOS coatings tend to have poor mechanical strength, limiting their reliability in practical applications. In this experimental study, we investigate the effect of adding a new functionalized FCOS to aluminized solid propellants. Specifically we focus on the oxidation and ignition characteristics of Al-FCOS composite particles and on the combustion of fluorine-containing propellants. By altering the surface pre-ignition reaction, FCOS-modified particles are found to ignite more quickly and produce more gaseous products. For pressures above 5 MPa, adding FCOS can both increase and decrease the burning rate relative to a baseline Al propellant, depending on the precise Al-FCOS content used, with 8.5 wt% and 17 wt% Al-FCOS producing a higher and lower burning rate, respectively. Compared with the baseline Al propellant, the FCOS-modified propellants are found to produce smaller agglomerates. This shows that the ignition, combustion and agglomeration characteristics can be readily tuned by adjusting the content of the new functionalized fluoropolymer. These findings can help guide the development of FCOS additives for aluminized propellants in solid rocket motors.
Combustion characteristics of a rapid mixed swirl torch igniter of CH4/O2 for hybrid rocket motors has been experimentally investigated. The igniter torch consists of four tangential slits ...circumferentially equispaced on the internal micro-combustion chamber allows tangential injection of O2 gas, meanwhile the CH4 flows into the main channel and interacts with four tangential O2 injections flow to generate a swirl mixing of CH4 and O2. The ignitability of this swirl torch igniter was verified by large number of experiments in variation of equivalence ratio and total flowrate of CH4/O2. The flame structure of the igniter and effect of main oxidizer injection acting as a cross flow on the flame has been investigated by using OH* and CH* chemiluminescences techniques. It is found that the torch igniter can reliably ignite instantaneously in the range of equivalence ratio 0.2∼1.4 and the effect of main oxidizer injection on the flame is tiny and can be neglected. In addition, the performance of this rapid mixed swirl flame has been investigated by using a lab scaled hybrid rocket motor and compared with the same hybrid rocket motor ignited by catalytic bed. The performance of this igniter and its comparison with catalytic bed ignition will be discussed and analyzed in detail.
The mass and thermal flow of micropolar fluid in a channel having permeable and moving walls perpendicular to the flow-direction is considered here. A mathematical formulation for this physical ...problem is made using the Eringen's micropolar fluid-thermal model. Closed form solutions for temperature, microspin of aciculate particle and velocity are derived using the double perturbation method followed by the similarity transformation for the Newtonian/micropolar fluids. The perturbation parameters are the Reynolds number (that controls the wall-injection/suction) and the wall dilation parameter (that controls the rate of flow through the pores). Usually the no-spin conditions are imposed on studying the micropolar fluid-flows, but we consider no-spin condition in this attempt in order to highlight effect of micropolar spin boundary layer in the vicinity of the walls of the channel. The convergence of the perturbation method is checked, that sounds well. The results obtained for particular case (Newtonian fluid) are compared with alike literature and they are found satisfactory. The results for velocity, temperature, shear stress, thermal transfer rate and spin are presented for different values of physical parameters quantitatively and qualitatively.
Solid rocket motors are complex systems which need to withstand extreme physical conditions in terms of temperature, pressure, and high-density energy release. Therefore, specific attention should be ...brought to the flaws that may occur during motor manufacturing\handling phases prior to launch. An example of such flaws is debonding, usually arising at the interface between case insulation and solid grain. When debonding is significant in size, it may result in the premature case exposure to combustion chamber hot gases, and, in worst cases, it may even cause a complete motor failure. This work is intended to evaluate the impact of propellant debonding on solid rocket motor case-insulating layer, making predictions about the most unfavorable regions where the debonding could occur. Numerical simulations are performed with an in-house simulation software applied to an actual solid rocket motor stage.
•An algorithm able to evaluate the effects on solid rocket motors grain debondings is proposed.•The present approach can estimate debondings effects on thermal protection case-insulating layer of solid rocket motors.•The main outcome is the identification of the most critical debonding positions on a solid rocket motor case.•Contour maps are carried out based on shape and size of the investigated debondings.
With the aim of enhance the designability of burning surface, star and wagon wheel fuel types are applied extensively in hybrid rocket motors. However, mechanism of nozzle ablation in hybrid rocket ...motors with these complex fuel types is unclear, and this increases the difficulty of predicting ablation laws and thrust. This paper intends to investigate spatial distribution characteristics of nozzle ablation in the hybrid rocket motor with various fuel types. According to the principle that the grain port area and grain length are equal, tube, star, single-port wagon wheel, and multi-port wagon wheel grains are designed. Numerical models for combustion flow and thermochemical erosion are established, and three-dimensional numerical simulations are conducted. A firing test for nozzle erosion is conducted, and hydroxyl-terminated polybutadiene and 95 % hydrogen peroxide are employed. Errors of throat erosion rate and combustion chamber pressure among simulation results and experimental data are 4 % and 4.25 %, respectively. Calculation results demonstrate that the circumferential distributions of nozzle wall temperature and erosion rate show an obvious correlation with fuel types. The maximum throat erosion rates for star, single-port wagon wheel, and multi-port wagon wheel grains occur at the circumferential position of 0° (star slot, spoke slot, and midpoint of outer edge of wheel hole), while the minimum throat erosion rates locate at the circumferential position of 36° (star root, spoke root, and midline of adjacent wheel holes). The circumferential non-uniformity of throat erosion rate of multi-port wagon wheel grain is the highest. When the grain port area and oxidizer mass flow rate are the same, the throat erosion rate of multi-port wagon wheel grain is the highest, followed by the star grain, and the throat erosion rate of single-port wagon wheel grain is the lowest.
•Effect of complex fuel types on spatial characteristics of nozzle erosion is studied.•Firing test of carbon-based nozzle erosion is performed to verify numerical models.•Three-dimensional simulations on erosion in hybrid rocket motors are conducted.•Simulation results agree well with test data.•Ablation mechanism of star, single-port, and multi-port wagon wheel grain is revealed.
Lunar missions are currently experiencing a significant surge in popularity, presenting expansive opportunities for further exploration and development. To thoroughly explore the design margins and ...potential of lunar landers, and to foster the development of overall designs driven by comprehensive performance objectives, it is crucial to conduct optimization design considering the coupling between key disciplines such as trajectory and propulsion. Considering the significant increase in computational complexity caused by conducting trajectory/propulsion integrated design optimization, the analytical target cascading method is employed to hierarchically decompose and coordinate optimization of the complex systems. This article presents a phased soft-landing strategy on the manned lunar lander propelled by hybrid rocket motors, utilizing powered explicit guidance and Apollo powered descent guidance, and proceeds with the trajectory/propulsion integrated design optimization involving diverse grain shapes and feed systems. This optimization process is separately undertaken utilizing multidisciplinary feasible method and analytical target cascading method. The analysis reveals that integrating trajectory and propulsion considerations into the optimization process facilitates a 5 % reduction in the overall mass relative to optimizations constrained solely by velocity increment and lack comprehensive trajectory design considerations. This highlights the profound impact of trajectory requirements on propulsion system design and the advantages of powered explicit guidance laws in minimizing fuel consumption. Crucially, the use of analytical target cascading achieves the better optimization results, and significantly reduces subsystem evaluation times, enhancing operational efficiency by 48 %, demonstrating the advantage in handling complex, large-scale systems. On another level, with different β values, the Mean Relative Error of the target values for the three schemes obtained by the analytical target cascading method is 0.0016, indicating good stability and strong robustness. The practical exploration in this article provides methods and frameworks for high-performance optimization design of complex aerospace mission profiles in the future.
•The trajectory/propulsion integrated design optimization is conducted.•The two-phase guidance strategy for lunar lander descent is proposed.•Based on the strategy, the analytical target cascading-decomposed framework is set.•The influence of trajectory and propulsion coupling on overall design is revealed.•The optimization effectiveness, convergence rate, and robustness of the ATC is shown.
Total impulse is one of the most important parameters which is highly related to the performance of solid rocket motors. Predicting the total impulse accurately is necessary for both design and ...operation purposes. However, the traditional methods greatly rely on expert knowledge and are less capable of analyzing modern sophisticated equipment. In this paper, a deep learning-based total impulse prediction method is proposed for the ignition process of solid rocket motor. A CNN-LSTM-Attention deep neural network model is established, which can automatically process raw data for feature extraction, efficiency improvement and prediction with high accuracy. Practical rocket data are used for validations which are collected in the ignition process. We compared the proposed method with the other popular algorithms to verify the effectiveness and superiority of this method. The results show that the proposed data processing and prediction method can achieve promising performance. The best result of average percentage error on the test set is below 2%. The dependency of the deep learning-based method on the data amount is largely reduced by using the downsampling method in data processing. In this way, the proposed method is available even with very few training data, which has good application prospects in engineering problems and provides an approach for combining artificial intelligence and solid rocket motor research.
This paper reports the development of a laser ignition system for a super-small solid rocket motor to be used in the OMOTENASHI, a 6U-CubeSat that is the smallest lunar lander ever. A safe, compact, ...lightweight, and optimized system using a laser diode was developed to ignite the rocket motor. The development was successful, and the spacecraft equipped with the laser ignition system was delivered to NASA. The laser ignition system will be the first one for a spacecraft going on a deep space exploration.
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