Propellant Combustion Process Research Articles

Experimental and numerical investigation on agglomeration of aluminum during combustion process of aluminized composite propellant

The addition of aluminum particles can effectively improve the energy characteristics of solid composite propellants in propulsion systems. However, these aluminum particles affecting propellant mixture viscosity, ignition delay time, burning time and specific impulse could also lead to undesirable effects. Therefore, it is a general practice that in the early stage of propellant formulation design that the particle size of condensed phase products is optimized/minimized through formulation adjustment to reduce its negative impact on the engine. In practice, it means that a costly manpower and materials are needed to obtain the size distribution of condensed phase products during the combustion process of propellants via experimental tests. Thus, there is a strong industrial need for a reasonable and feasible simulation method of agglomeration process during the combustion process of propellants. In this work, we propose and examine the implementation of molecular dynamics and a random packing algorithm on creating numerically models of aluminized composite propellants. Considering the pyrolysis of ammonium perchlorate (AP) and the combustion process of aluminum, we develop a more accurate agglomeration model basing on the classical Jackson model. Our proposed methods and simulations models are validated by our experimental results. It confirms that the present work provides a prediction tool for simulating and improving the agglomeration process, thus aiming at optimizing solid propellant formula.

Elaborative collection of condensed combustion products of solid propellants: Towards a real Solid Rocket Motor (SRM) operational environment

A novel constant-pressure and constant-quenching distance Condensed Combustion Products (CCPs) collection system was developed, coupled with a timing control system, to collect the CCPs formed in the course of burning of aluminum-based composite propellants. The effects of adiabatic graphite plating, collection zone, quenching distance, time series of collection, and propellant burning rate on the microscopic morphology, particle size distribution and unburned aluminum content of CCPs were investigated. It was verified that the graphite plating can provide a high-fidelity high-temperature environment for propellant combustion. The combustion efficiency is improved by 2.44% compared to the bare propellant case. The time series of collection has a significant effect on the combustion efficiency of aluminum, and the combustion efficiency of aluminum in the thermal state (1.2–2.4 s) is 2.75% higher than that in the cold state (0–1.2 s). Similarly, the characteristics of the CCPs in different collection zones are different. At the quenching distance of 5 mm, the combustion efficiency of aluminum in the core zone (85.39%) is much lower than that in the outer zone (92.07%), while the particle size of the CCPs in the core zone (172 μm) is larger than that in the outer zone (41 μm). This indicates that the core zone is more likely to produce large-sized and incompletely burned agglomerates during the propellant combustion process. Different burning rates also lead to a significant difference in particle size distribution and combustion efficiency. High burning rates result in higher combustion efficiency. A detailed sequence of the elaborative collection process of CCPs is proposed, mainly including the setting of ignition delay time, burning rate, working pressure, plating length and time series of collection. The findings of this study are expected to provide a reliable tool for the evaluation of the combustion efficiency of solid propellants.

Estimation of surface geometry on combustion characteristics of AP/HTPB propellant under rapid depressurization

The 2D sandwich model serves as a potent tool in exploring the influence of surface geometry on the combustion attributes of Ammonium perchlorate/Hydroxyl-terminated polybutadiene (AP/HTPB) propellant under rapid pressure decay. The thickness of the sandwich propellant is derived from slicing the 3D random particle packing, an approach that enables a more effective examination of the micro-flame structure. Comparative analysis of the predicted burning characteristics has been performed with experimental studies. The findings demonstrate a reasonable agreement, thereby validating the precision and soundness of the model. Based on the typical rapid depressurization environment of solid rocket motor (initial combustion pressure is 3 MPa and the maximum depressurization rate is 1000 MPa/s). A-type (a flatter surface), B-type (AP recesses from the combustion surface), and C-type (AP protrudes from the combustion surface) propellant combustion processes are numerically simulated. Upon comparison of the evolution of gas-phase flame between 0.1 and 1 ms, it is discerned that the flame strength and form created by the three sandwich models differ significantly at the beginning stage of depressurization, with the flame structures gradually becoming harmonized over time. Conclusions are drawn by comparison extinction times: the surface geometry plays a pivotal role in the combustion process, with AP protrusion favoring combustion the most.

Influence of pre-ignition technology on combustion and explosion characteristics of composite propellants

The slow cook-off test is widely used for analyzing the safety of propellants under accidental thermal stimulation. To examine the influence of the pre-ignition technology of a composite propellant on thermal safety and combustion characteristics, a quantitative evaluation experiment of slow cook-off reaction intensity of a typical composite propellant charge was designed and executed, considering hydroxyl-terminated polybutadiene (HTPB) and hydroxyl‑terminated block copolyether (HTPE) propellants containing ammonium perchlorate (AP) and aluminum (Al). The experimental method for quantitative evaluation of the slow cook-off reaction intensity is highly essential to investigate the safety of ammunition. The influence of pre-ignition on reaction intensity was analyzed at different temperatures, the combustion behavior of the heated propellant was studied using a closed bomb experiment, and the influence of pre-ignition on combustion evolution was analyzed. Finally, the process was characterized by theoretical modeling. The results show that the slider speed was 169.3 m/s when the HTPB/AP/Al propellant was ignited at 234 °C. The slider speed was reduced to 43.4 m/s and 61.8 m/s by pre-ignition at 120 °C and 214 °C, respectively. When the spontaneous ignition temperature of HTPE/AP/Al propellant was 212 °C, the slider speed was 110.5 m/s. Pre-ignition at 120 °C and 192 °C reduced the slider speed to 39.4 m/s and 43.7 m/s, respectively. Pre-ignition can effectively reduce the reaction intensity. Under the same conditions, the reaction strength of HTPE/AP/Al propellant is marginally lower than that of HTPB/AP/Al propellant. In the slow heating process before ignition, with the increase in ignition temperature, when HTPB/AP/Al propellant was pre-ignition at 120 °C and 214 °C, the pressurization time was 9.8 ms and 6.4 ms, respectively, while before ignition, the pressurization time decreased to 5.5 ms. At 120 °C and 192 °C, the pressurization time of HTPE/AP/Al propellant was 50.9 ms and 36.6 ms, respectively, while before ignition, the pressurization time decreased to 26.9 ms. The reduction in ignition temperature decreased the pressure growth rate of propellant during combustion and reduced the pressurization time. A calculation model of the propellant combustion process was established to characterize the growth process of the combustion pressure at different ignition temperatures. This model can effectively predict the combustion evolution of composite propellants after the slow cook-off ignition and aid in ensuring propellant safety.

An Analysis and Calculation Method for One‐Dimensional Balanced Interior Ballistics of a Recoilless Gun

AbstractThe balancing force of the interior ballistics is an important issue in the design of the recoilless gun. It has been found in practice that the balancing force design method for recoilless guns based on the traditional classical interior ballistic theory cannot guide the balance from the physical mechanism and process. In order to study the unbalanced force experienced by the recoilless gun during the firing of the projectile and the variation law of its various parameters with time, the analysis and calculation method of the balanced interior ballistics is proposed in this article. This method separately models the aerodynamic and propellant combustion processes of the recoilless gun, using the characteristic line method to solve the model numerically. This method can be used to predict the change of the unbalanced force acting on the recoilless gun during the projectile firing of the recoilless gun. The reliability of the algorithm is verified by the interior ballistic performance when China 78 type 82 mm recoilless smoothbore gun launches the projectile. The study found that the unbalance impulse generated by the gas inside the barrel fired by the recoilless gun accounts for only 1.5 % of the total recoil impulse, but it has a great influence on the amount of impulse a shooter has can take. This approach provides a better understanding of this complex internal ballistic process, which has important implications for the design and improvements of recoilless guns.

DESIGN OF A TRANSMISSION SYSTEM ON A BALL MILL MACHINE TO PROCESS HOMOGENIZATION OF BLACK POWDER USING AN AUTOMATIC SYSTEM

Abstract: In the military world, battle technology has experienced very rapid development, both close-range combat and long-range combat. In long-distance combat the technology that is often used is rocket technology where this rocket technology can paralyze the enemy in large numbers. Black powder is a mixture of potassium nitrate (KNO3), carbon powder (C) and sulfur (S). (Evie Lestariana, LAPAN). The black powder itself functions as an igniter filling material for the propellant combustion process. This research is to design a tool to support the manufacture of black powder in the igniter filling to help the combustion of rockets, especially those owned by the Indonesian Army. This study uses an experimental method using empirical calculations to get a tool with the desired specifications. Before carrying out the calculations, the author collects research data to carry out calculations, the data collected are as follows: 1. Roller shaft rotation: 246 rpm, 2. Roller shaft torque: 4.0697 Nm, 3. Roller shaft diameter: 25 mm, 4. Pulley shaft diameter top: 100mm. The results obtained after carrying out the calculations are the diameter of the driving pulley 98 mm, the belt contact angle against the pulley 174.6º, and the length of the V-belt 453.14 mm, moment of inertia on the gearbox output shaft 1884,8 mm4.

Characterization of Aluminum Powders: IV. Effect of Nanometals on the Combustion of Aluminized Ammonium Nitrate‐Based Solid Propellants

AbstractThis study is devoted to the investigation of the energetic properties of two types of aluminized solid propellants with inert and active binders modified by additives of six different metal nanopowders. The substitution of an inert binder for the active one resulted in significant improvement of energetic properties. However, burning rate regulation become more complicated because the pressure exponent in burning rate law was increased as well. The DSC analysis revealed that the most significant catalytic effects were observed while using Zn, Cu, and Mo powders. Combustion of compositions based on an inert combustible binder and ammonium nitrate was studied in the pressure range 0.1–10.0 MPa. Due to the low combustion rate, no dispersion of burning aluminum particles was observed. During combustion, a single agglomerate of aluminum particles was formed, the mass of which was about 96 % of the initial aluminum mass in the formulation. A comparative analysis of nanosized metal additive effect on the propellant's combustion process was performed. It had been found that the activity of the metal nanopowders depends on the propellant formulation. It had been shown that Al and Zn did not affect the combustion rate of propellant with inert binder but they were effective in compositions with an active binder. Cu and Ni were found to be a good catalyst for the ammonium nitrate and active binder decomposition.

Modeling and attitude control of CubeSat utilizing moving mass actuators

Any vehicle propelled by solid rocket motors (SRMs) must include an attitude control system capable of dealing with the torque generated by thrust misalignment. In order to expand the application of SRMs on CubeSats, an attitude control system utilizing moving mass actuators is discussed. The present research develops an eight-degree-of-freedom simulation model of a 2U CubeSat with two moving mass actuators. That model also considers the influence of propellant combustion processes. By analyzing the model disturbance source and systematic coupling, the key layout parameters are designed and a simplified control model is proposed. The controller is derived based on a combination of backstepping and sliding mode techniques. An orbit maneuver from 300 km circular orbit to 300 and 500 km elliptical orbit using this attitude control system is verified.

Thermal decomposition and combustion characteristics of Al/AP/HTPB propellant

Aluminum(Al)/ammonium perchlorate(AP)/hydroxyl-terminated polybutadiene (HTPB) propellant is the most widely used propulsion system currently. Its ignition and combustion processes are fairly complex and need to be better understood. In this study, the thermal decomposition, ignition, and combustion properties of an Al/AP/HTPB propellant were investigated using a thermogravimetry–differential scanning calorimetry and a laser ignition testing system. The morphology and size distribution of the condensed combustion products (CCPs) were analyzed using a laser particle size analyzer and scanning electron microscopy. Results showed that the thermal decomposition process of the propellant consisted of three stages. The first stage (100–430 °C) was a major mass loss stage and exhibited typical features of AP decomposition. The second stage (430–630 °C) was mainly accompanied by the decomposition of remaining HTPB as well as slight oxidation of Al particles. In the third stage, further oxidation of Al particles resulted in a small mass increase. Due to the continuous emission problem, only a few combustion intermediates were identified in the combustion emission spectrum curves. The propellant combustion process could be roughly divided into three stages, and the flame development stage lasted longer than the flame decline stage. As the pressure increased, the propellant ignition delay time decreased and the burning rate increased significantly. The effect of pressure on ignition delay was more pronounced at low pressures. The CCPs consisted of three types. The oxidation of Al particles in the propellant followed the diffusion reaction mechanism. Agglomerates appeared in several different morphologies and had relatively low combustion efficiency under experimental conditions. A 2-D pocket-like structure was observed on the propellant surface.

Damage localization effects of the regeneratively-cooled thrust chamber wall in LOX/methane rocket engines

To investigate the damage localization effects of the thrust chamber wall caused by combustions in LOX/methane rocket engines, a fluid-structural coupling computational methodology with a multi-channel model is developed to obtain 3-demensioanl thermal and structural responses. Heat and mechanical loads are calculated by a validated finite volume fluid-thermal coupling numerical method considering non-premixed combustion processes of propellants. The methodology is subsequently performed on an LOX/methane thrust chamber under cyclic operation. Results show that the heat loads of the thrust chamber wall are apparently non-uniform in the circumferential direction. There are noticeable disparities between different cooling channels in terms of temperature and strain distributions at the end of the hot run phase, which in turn leads to different temperature ranges, strain ranges, and residual strains during one cycle. With the work cycle proceeding, the circumferential localization effect of the residual strain would be significantly enhanced. A post-processing damage analysis reveals that the low-cycle fatigue damage accumulated in each cycle is almost unchanged, while the quasi static damage accumulated in a considered cycle declines until stabilized after several cycles. The maximum discrepancy of the predicted lives between different cooling channels is about 30%.

Investigation of Stick Propellant Internal Perforation Erosive Burning on Interior Ballistics Performances

In this study, a thorough investigation of a stick propellant internal perforation erosive burning on interior ballistic performances is presented via extending the previous work of author. The stick propellant combustion process and the internal perforation erosive burning are revealed by numerical simulations. Different factors with respect to the stick propellant, including propellant length, internal perforation diameter and loading density are analysed in detail. Stick propellant length and the internal perforation diameter have a significant influence on the ballistic performance, the longer and smaller internal propellant have a greater erosive burning effect to the ballistic performance. Loading density has very weak influence on the ballistic.

Observation on Solid Propellant Ignition Using High Speed Camera: The Effect of Hot Wire Position on the Combustion Flame Contour

This paper reports on the discovery of unique flame structure of a composite propellant sample under hot wire ignition. The entire combustion process at atmospheric pressure condition was recorded using a high speed camera. Three hot wire orientations were chosen in this experiment for examining their effects on the propellant burning behavior. The results show that the wire orientations are crucial in propellant combustion process, as different flame patterns were observed when the hot wire orientation was altered. This paper provides an important insight into this specific ignition process that can be useful for researchers in the aerospace industry for better design and more realistic simulation results in ignition control.

Experimental Studies of Propellant Loading Parameters and Plasma Flow-Field Interactions

Propellant-plasma interaction has been attracting great interests among most researchers for application to an electrothermal-chemical (ETC) gun. Previous works have shown that propellant-plasma interaction can enhance burning rates due to increased surface erosion of propellants as compared to that of conventional ignition. Plasma radiation and heat convection are the dominant mechanism of energy transfer based on both experimental data and numerical simulation. However, little literatures were reported about a detailed understanding of propellant loading parameters and the plasma flow-field distribution interactions, which is very important in deciding a successful propellant-charge design in ETC guns. The pressure-time curves during interrupted-burning process of different propellants such as triethyeneglycol dinitrate and single-base propellants ignited by plasma were recorded by pressure sensor in interrupted-burning setup. Influence of loading parameters, including loading density, distance between propellants and plasma jet, and propellant holding methods at the same selected location in plasma flow-field on plasma-ignition performance of propellants was discussed based on experimental results. As the loading density of propellants increases, the porosity of the propellant bed becomes small. Moreover, the penetrating depth of plasma into propellant bed shortens. Therefore, the plasma energy attenuates rapidly, and the ignition-delay time of propellants becomes longer. The ignition-delay time is longer when the distance between propellants and plasma jet increases. At the same selected location in plasma flow-field, different holding methods of propellant burning surface to plasma jet have less influence on the ignition, combustion process of propellants, and the ignition-delay time. The matching test results on propellant loading parameters and plasma flow-field distribution interaction play a major role in improving charge configurations at chosen plasma flow-field and understanding plasma-propellant interaction better.

A MEMS-based solid propellant microthruster with Au/Ti igniter

A solid propellant microthruster with Au/Ti igniter is demonstrated as an improved micropropulsion system for microspacecraft. The new design provides the microthruster with a high degree of flexibility, maneuverability and integration. Single microthruster and microthruster arrays have been successfully fabricated using standard microfabrication technologies. The propellant combustion process in the micro-chamber of the microthruster has been visually observed. Initial tests employing gunpowder-based solid propellants, have produced 2.11 × 10 −5 to 1.15 × 10 −4 N s of total impulse and 2.68–14.65 s of specific impulse at sea level, and 3.52 × 10 −5 to 2.22 × 10 −4 N s of total impulse and 4.48–28.29 s of specific impulse in vacuum. The performance of the solid propellant microthruster with Au/Ti igniter is also compared with that of a solid propellant microthruster having a wire igniter.

Catalyzed combustion of hydrogen–oxygen in platinum tubes for micro-propulsion applications

The present investigation addresses the need to understand the physics and chemistry involved in propellant combustion processes in micro-scale combustors for propulsion systems on micro-spacecraft. These spacecraft are planned to have a mass less than 50 kg with attitude control estimated to be in the 1–10 mN thrust class. Micro-propulsion devices behave differently than macro-scale devices because of the differences in magnitude of flow rates and heat transfer. Reducing the combustor size increases the relative surface area, increasing the heat loss, and as combustors are continuously reduced in size, they approach the quenching dimensions of the propellants. Combustors of this size are expected to significantly benefit from surface catalysis processes. A miniature flame tube apparatus is chosen for this study because microtubes can be easily fabricated from known catalyst materials, and their simplicity in geometry can be used in fundamental simulations for validation purposes. Experimentally, we investigated the role of catalytically active surfaces within 0.4 and 0.8 mm internal diameter microtubes, with special emphases on ignition processes in fuel rich gaseous hydrogen and gaseous oxygen. Calculations of flame thickness and reaction zone thickness predict that the diameters of our test apparatus are below the quenching diameter of the propellants in most atmospheric test conditions. The temperature and pressure rise in resistively heated platinum microtubes and the exit hydrogen concentration were used as an indication of exothermic reactions. Data on imposed heat flux/preheat temperature required to achieve ignition versus mass flow rate are presented. With a plug flow model, the experimental conditions were simulated with detailed gas-phase chemistry and surface kinetics. Computational results, in general, support the experimental findings.

Spectral Acquisition and Calibration Techniques for the Measurement of Radiative Flux Incident upon Propellant

AbstractA technique for the absolute calibration of a time‐resolved spectrographic system has been developed at QinetiQ, specifically designed to be relevant to spectral acquisition from within the interior of translucent gun‐propellant samples. The technique has shown itself to be particularly useful in the realm of propellant ignition as it allows for the precise determination of the moment that propellant combustion processes begin, as well as measuring the incident radiative flux leading up to ignition. Scope exists to extend its use for high‐pressure measurements of the incident radiative flux during both conventional propellant burn and high‐powered electrothermal‐chemical (ETC) discharges. This paper sets out to describe both the technique and some of the pitfalls encountered during the development of the technique. The use of this technique in some of the experimental work performed at QinetiQ, including the results of measurements that compare the incident radiative flux with propellant ignition during both ETC discharges and conventional gunpowder burn, have been published separately; references for this experimental work are given in this paper.

Experimental characterization of plasma effects on energetic materials for electrothermal-chemical launch applications

Energetic electrical plasmas play a fundamental role in the electrothermal-chemical (ETC) propulsion concept, where they are used to provide energy conversion from an electrical power source to the combustion chamber of an ETC launcher. Recent research in the areas of large-caliber ETC gun experiments as well as small-scale plasma experiments have demonstrated the ability to alter the propellant combustion process and the chemical or physical properties of propellant samples exposed to energetic electrical plasmas. Such results have warranted continued investigations into the underlying processes involved in propellant and plasma interactions, specifically in the behavior of exposed propellant samples to plasmas operating at varying degrees of electrical power. Experimental characterization has been performed with various samples of energetic materials including JA2, CL20, RDX, M9, M30, M43, KELP, and BAMO-AMMO (BA), which were allowed to interact with plasmas having different electrical power levels. The plasma power level ranged from 60 MW to a maximum of 264 MW, and total energy transferred to the plasma capillary ranged from 17 to 70 kJ. Propellant samples were analyzed using a Fourier transform infrared spectroscopy (FTIR) technique, which was used to determine the effects of plasma exposure on the energetic materials. Preliminary results from the experiments indicate that some propellant formulations are easily altered (either chemically or physically) or ignited by a plasma while others are less sensitive to the plasma exposure. In addition, the effects of the plasma are found to be minimized with the introduction of protective films even for plasmas of high power and energy levels.

Combustion behaviour of composite propellant at sub-atmospheric pressure

A family of four composite propellants has been investigated with regard to its combustion behaviour at low pressure. The effect of initial propellant temperature is very severe and sometimes leads to unexpected results, the burning rate being higher at low temperatures. The propellant temperature sensitivity σ p shows a correlation with temperature and propellant composition. Combustion instability predominantly occurs below 55 kPa and seems to be related to the propellant combustion process itself.

Combustion Processes of Propellants with Embedded Metal Wires

Research was made with double-base propellants with embedded metal wires to understand how the burning rate changes along the silver wire and to obtain what factors control the burning rate of the propellants. The burning rate along the wire is influenced greatly by the kind of metal wire, wire size, and propellant compositions. With double-base propellants having varied burning characteristics, burning rate along the silver wire was measured and the effects of the luminous flame, dark zone, and the fizz zone were studied. By means of microthermocouples, the temperature profiles in the propellant near the silver wire were measured. Based on the test results, it was clear that the factors which control the burning rate along the silver wire are the temperature gradient in the fizz zone and the temperature in the dark zone. The luminous flame zone that stands above the burning surface affects neither the burning rate of the propellant nor the burning rate along the silver wire.

Solid propellant combustion in the presence of photoirradiation

Previous studies of the nonsteady processes associated with the irradiation of propellants with light have chiefly been devoted to the question of ignition [1–3]. It is also important to consider the effect of such an easily controlled influence as light on the propellant combustion process. We have attempted to estimate the dependence of the propellant burning rate on the intensity of the luminous radiation. Cases of steady-state combustion and combustion in the presence of a light flux varying harmonically with time are considered. It is assumed that the incident light flux is absorbed in the solid phase in accordance with the Bouguer-Lambert exponential law with constant transparency index. Steady-state combustion is considered within the framework of the Zel'dovich theory [4]. It is shown that in the steady state irradiation is equivalent to a certain increase in the initial temperature of the propellant. In the case of combustion with irradiation this makes it possible to use the data on steady-state combustion without irradiation. Nonsteady combustion in the presence of a periodically varying light flux is described with the aid of the Novozhilov model [5]. A correction to the mean burning rate (Δu °), proportional to the square of the light flux amplitude, is obtained. In the case of an exponential dependence of the burning rate on initial temperature the correction Δu ° is negative. The effect of irradiation on the stability of the steady-state propellant combustion mode is discussed.