Conventional Propellant Research Articles

Theoretical and Experimental Analysis of the Cathode-Vapour-Feed PEM-Electrolyser for Space Applications

Water electrolysis is experiencing growing interest due to its importance in the transition to a carbon-neutral transportation sector and even in spaceflight it can play a vital role. Here, its typical application has been the oxygen generation for the life support of astronauts onboard a spacecraft or the International Space Station. However, in recent years the application of electrolysis for a new form of spacecraft propulsion is receiving increasing attention, as well. This technology is called Water Electrolysis Propulsion (WEP).So far highly toxic and expensive propellants have been used for in-space propulsion, like hydrazine-derivatives and dinitrogen tetroxide. However, with a constantly growing number of satellite-launches per year, the spaceflight industry is searching for more inexpensive solutions which do not entail the corresponding safety hazards and engineering challenges of these conventional propellants. WEP is currently considered to be one of the most promising solutions for these issues. Within such a system a satellite is filled on ground with pure water. Once launched into space an electrolyser, powered by the satellite’s solar panels, is used to split the water into gaseous hydrogen and oxygen which are stored in intermediate storage tanks at pressures of up to 100 bar. Subsequently the gases can be combusted in a rocket engine to generate thrust and to propel the spacecraft for various propulsive needs.One of the core components of such a system is the electrolyser. In order to be competitive, an electrolyser type has to be found, which is lightweight, able to operate in zero-gravity and is able to pressurize the gases without the need for additional mechanical pumps. Currently the most promising electrolyser type is the so-called Cathode-Vapour-Feed (CVF) PEM electrolyser. Here a conventional PEM electrolyser is operated in cathode feed and modified by integrating a second membrane (Water Feed Barrier) between the electrolysis membrane and the water inlet. In order for the electrolyser to operate, the water has to diffuse through this Water Feed Barrier (WFB) and is subsequently present in a vapour state at the electrolysis membrane. Therefore, no phase separators are needed and the electrolyser is able to operate independently and unaffected by gravity. Furthermore, a low voltage is applied on the Water Feed Barrier slightly exceeding the Nernst-Voltage. Hence the WFB is effectively acting as an electrochemical pump which allows the generation of the gases at higher pressures than the water inlet pressure.This conceptual change is however introducing several unconventional mechanisms and phenomena that have to be considered during the design of such a device. However, previous research on the CVF technology has been very limited due to its niche application so far and most of these phenomena and mechanisms remain unanalysed. This paper is aiming to contribute towards the closure of this knowledge gap.The working principle and the theory behind these additional phenomena appearing in the CVF electrolyser are presented. These are the significantly affected mass transport of the water towards the anode side of the electrolysis membrane, the gas diffusion through both membranes and its effect on the electrolyser’s performance, as well as the mutual interaction between the applied voltages on the three electrodes. Furthermore, many researchers have used the same membrane type for the WFB as for the electrolysis membrane, although it serves a different purpose and is exposed to different operating conditions. Therefore, special attention will be devoted to the determination of the optimal membrane type for the WFB since barely any research has been conducted on this topic. In addition, a proof-of-concept electrolyser has been built and tested in a parameter study to validate the theoretical considerations. The experimental findings on the effect of membrane selection for the WFB, distance between the membranes and impact of flow field topology are presented and discussed. Therefore, the paper contributes towards a more targeted development of space electrolysers in the future. Figure 1

Optical-Propulsion Metastructures.

Pulsed laser micropropulsion (PLMP) offers a promising avenue for miniature space craft, yet conventional propellants face challenges in balancing efficiency and stability. An optical-propulsion metastructure strategy using metal-organic frameworks (MOFs) is presented to generate graphene-metal metastructures (GMM), specifically GMM-(HKUST-1), which significantly enhances PLMP performance. This novel approach leverages the unique interaction between pulsed lasers and the precisely engineered GMMs-comprising optimized metal nanoparticle size, graphene layers, and inter-particle gaps-to boost both propulsion efficiency and stability. Experimental and numerical analyses reveal that GMM-(HKUST-1) achieves aspecific impulse of 1072.94 s, ablation efficiency of 51.22%, and impulse thrust per mass of 105.15 µN µg-1, surpassing traditional propellants. With an average particle size of ≈12 nm and a density of 0.958gcm-3, these metastructures exhibit 99% light absorption efficiency and maintain stability under atmospheric and humid conditions. The graphene nanolayer efficiently absorbs and converts laser energy, while the metal nanostructures enhance light-matter interactions, promoting energy transfer and material stability. These findings suggest that this GMM-based optical-propulsion strategy can revolutionize microspacecraft propulsion and energy systems, offering significant advancements across various domains.

Evaluation of Boron Combustion for Ducted Rocket Applications Using Condensed Product Analysis

Boron, a metalloid, produces high energy upon combustion. It is recommended as an ingredient for fuel or propellant in rocket propulsion, despite the challenge of extracting its full thermal energy. So far no one has claimed the complete energy conversion of boron upon combustion. On the other hand, the current propulsion system of the Meteor missile uses boron-loaded propellant. The boron-loaded propellant provides an approximately three-fold increase in specific impulse compared to conventional propellants. The present study focuses on boron-HTPB-based solid fuels impregnated with early ignited particles as additives, aiming to assess the combustion performance of boron particles. These additives are magnesium (Mg), titanium (Ti), and activated charcoal (C), and their effects are evaluated based on the residual active boron content in the condensed combustion products (burned residues). An economical tool commonly called stagnation flow or opposed flow burner (OFB) is used to deflagrate the fuel sample by means of pressurized oxygen gas. The condensed combustion products are examined using a field emission scanning electron microscope (FESEM), X-ray diffraction (XRD), thermogravimetric analysis (TGA), and differential thermal analysis (DTA). Among the fuel combinations investigated here, magnesium has been found to be a good burning enhancer of boron, leaving the lowest active boron content (30%) in the burned-out residue. The current research aims to develop an efficient boron-containing solid fuel for hybrid propellant ducted rocket engine applications.

MOFs for Ultrahigh Efficiency Pulsed Laser Micropropulsion.

Conventional propellant materials, such as polymers and single metal elements, have long been investigated for their potential in pulsed laser micropropulsion (LMP) technology. However, achieving superior LMP efficiency through physical mixing of these materials remains a significant challenge. This study presents a paradigm shift by introducing porous crystalline polymers, known as metal-organic frameworks (MOFs), as novel propellants in pulsed LMP. MOFs are composed of metal cations and organic ligands that form ordered structures through coordination, eliminating the problem of local hot zones arising from uneven physical mixing encountered in LMP. In direct comparison to conventional polymers and single element targets, MOFs exhibit substantially higher LMP efficiency. By precisely tailoring the metal atom fraction within MOFs, an extraordinary ultrahigh efficiency of 51.15% is achieved in pulsed LMP, surpassing the performance of similar materials previously reported in the literature. This pioneering application of MOFs not only revolutionizes the field of LMP but also opens up new frontiers for MOF utilization in various energy applications.

Development of high burn rate propellant and testing in miniature rocket motor for control applications

For aerospace vehicles, notably missiles or satellites for attitude and divert control, impulsive thrusters are most desirable if they are quantized as exact instantaneous pulses available over and after extended periods of inactive storage. These thrusters operate for a few milliseconds (ms), typically between 10 ms to 100 ms duration. Compared to its liquid or gaseous counterpart, an impulsive thruster powered by solid propellants is a simple, and reliable solution. Therefore, they are suitable as reaction control systems and divert thrusters for satellite and unmanned aerial vehicle applications.A typical solid propellant would need an extremely thin web thickness for port burning arrangement due to its brief burn duration, which makes propellant manufacturing and its structural anchoring in motor challenging. A high burn rate and high-density propellant based on ammonium perchlorate, aluminum, and Teflon is processed to enable end-burning configuration thus solving the problems, simplifying propellant manufacturing, structural design and integration. It also permits stacking propellants for multi-pulse applications. The high burn rate (40 mm/s at 120 bar pressure) propellant is formulated and tested for ballistic performance evaluation in proof motor configuration, and evaluation of its advantages over conventional propellant. A case of identical design requirements with both double base propellant and high burn rate pressed propellant is compared, tested and evaluated to compare the relative performance merits. The volumetric load is increased from 47% to 82% due to the development and adoption of high burn rate propellant. The overall weight of the thruster is also reduced by 20%.

Study of Energetic Materials of Double Base Propellant for Small Caliber Ammunition

Double Base Propellants (DBP) has been frequently utilized in the fields of rocket propellant and gun propellant. DBP has various advantages over conventional propellants, including easy formulation, high energy, and smokelessness. The research conducted in literature studies was applied by referring to the various energetic materials for DBP. The composition of a DBP consists of energetic raw material of Nitrocellulose (NC) and Nitroglycerin (NG). According to data analysis, NC can be manufactured from the raw material Nata de Coco, which has a nitrogen content of more than 12,5% in NC, which is in line with the levels of conventional military and propellant raw materials. Meanwhile, SMX (1,4-dinitrate-2,3-dinitro-2,3bis (nitratomethylene) butane) or GAP (Glycidyl Azide Polymer) on double-base propellants is suggested as potential alternative energy sources that can be utilized to substitute NG. The choice of new energetic material is influenced by the fact that Nata de Coco is abundant in Indonesia, while SMX and GAP in terms of manufacturing and handling are safety. The several parameters to characterize double-base propellants include thermal analysis, microstructure, surface area, burning rate value, and gaseous product. The results of thermal properties by DSC analysis on a DBP based on NC+NG+TiO2-trifluoroacrylate showed a change in exothermic decomposition at a temperature of 199.2oC. Meanwhile, TGA analysis of a DBP based on NC+NG+GAP showed mass decomposition at temperature 199,3oC. Furthermore, it was found that thermoplastic elastomers GAP have a significant effect on increasing the low-temperature mechanical properties of double-base propellants. As a result, the literature studies approach, which includes consideration of alternative energetic raw materials and characterization for double-base propellants, is the recommendation for obtaining superior performance double-base propellants with energetic materials.

Conceptual Design of Small-Sized Thruster Using Laser Ignition of High-Energy Monopropellant

Ammonium dinitramide (ADN)-based energetic ionic liquid propellants (ADN-EILPs) present considerable potential as monopropellants due to high energy density, high thermal/chemical stability, and low toxicity. Consequently, a chemical thruster system based on ADN-EILPs can be employed in the development of high-performance ultrasmall/small satellites. However, the characteristics of the ADN-EILPs present various problems in their ignition during the thruster development. To solve this problem, the authors have previously achieved the continuous-wave (CW) laser ignition of ADN-EILPs using laser absorbers composed of carbon wools. However, the conventional propellant injection of ADN-EILPs with carbon wool using high-pressure gas faces several limitations. Therefore, a propellant feed system suited for the ignition method is proposed here, as well as a conceptual model of a 0.5 U/0.5 N-class thruster operated by the CW laser ignition of ADN-EILPs (). Additionally, an attempt is made to manufacture a laboratory model (LM) thruster, and its fundamental operation properties are determined. The future research implications of this study include the further observation of the combustion of the decomposed gas before its emission from the throat of the LM thruster, along with the further development of the propellant feeding system proposed in this study and the hot-fire tests performed for the feeding systems.

Experimental Studies on Parametric Effects and Reaction Mechanisms in Electrolytic Decomposition and Ignition of HAN Solutions.

The green propellant hydroxylammonium nitrate (HAN) is a good alternative to the conventional propellants in space propulsion applications because of its low toxicity and high energy density. Electrolytic decomposition and ignition of HAN solution, an ionic liquid, is a promising approach. In this work, comprehensive experimental studies were conducted to examine effects of different electrolytic voltages, electrode surface areas, and HAN concentrations on the decomposition process. In the test cases, an optimum electrolytic voltage appears to exist, which leads to the fastest decomposition process. As the voltage increases, a larger electrode surface area on the anode side should be used to overcome an anodic inhibition phenomenon and accelerate the electrolytic process. A high concentration of HAN solution is preferred for its decomposition and ignition. Results also reveal that the electrolytic process of a HAN solution could eventually trigger thermal decomposition reactions, raising the maximum temperature to around 550 K at the final stage. A detailed chemical reaction mechanism was proposed, based on the experimental data and FTIR spectra analyses. Results obtained herein would provide fundamental understandings on the complex electrochemical and physical processes and should be helpful for future applications of the electrolytic decomposition and ignition technology.

Qualitative Analysis of Spray Characteristics of Impinging Jets Using A Gelled Non-Newtonian Propellant Simulant

The utilization of liquid and solid fuels for propulsion and combustion processes with Newtonian characteristics are widely known. However, recent studies are considering the application of shear-thinning non-Newtonian fuels as alternative simulant since they have some advantages compared to the conventional propellants although there are challenges of providing better spray performance during the corresponding process. This paper, therefore, presents the results of the experimental investigations of the “near-field” spray characteristics with utilizing imaging techniques which evaluates the sheet formation and breakup length of four different spray patterns produced by the jet impingement of a gelled non-Newtonian propellant simulant. The qualitative analysis of this study shows that the spray patterns are different compared to those that are shaped by using Newtonian liquid fuels. This could lead to supposition that the non-Newtonian rheology of the gelled propellant simulant postponed the sheet and ligaments breakup, including the mode change of the atomization. In addition, the atomization of the sheet at different flow parameters could occur due to the formation and the sheet wave instability when aerodynamic and hydrodynamic of their origin are closely considered. This was further supported by the occurrence of perforations in the sheet.

Preparation of Bi2O3/Al Core-Shell Energetic Composite by Two-Step Ball Milling Method and Its Application in Solid Propellant.

In this article, Bi2O3/Al high-density energetic composites with a core-shell structure were prepared by a two-step ball milling method using a common planetary ball milling instrument, and their morphology, structure, and properties were characterized in detail. Through a reasonable ratio design and optimization of the ball milling conditions, the density of the Bi2O3/Al core-shell energetic composite is increased by about 11.3% compared to that of the physical mixed sample under the same conditions. The DSC (Differential Scanning Calorimetry) test also showed that the exothermic quantity of the thermite reaction of the energetic composite reached 2112.21 J/g, which is very close to the theoretical exothermic quantity. The effect of Bi2O3/Al core-shell energetic composite on the energy performance of insensitive HTPE propellant was further studied. The theoretical calculation results showed that replacing the partial Al with Bi2O3/Al core-shell energetic composite can make the density of propellant reach 2.056 g/cm3, and the density specific impulse reach 502.3 s·g/cm3, which is significantly higher than the density and density specific impulse of the conventional composite solid propellant. The thermal test showed that the explosive heat of the HTPE (Hydroxyl terminated polyether) propellant also increased with the increase of the content of Bi2O3/Al core-shell energetic composite.

Recent developments in ceramic microthrusters and the potential applications with green propellants: a review

Conventional chemical propellants such as hydrazine and ammonium perchlorate have been used within the realm of contemporary space propulsion devices and are well established owing to their rich heritage. However, their limitations such as toxicity, difficulty in operational handling and environmental impacts have raised concerns. In view of these limitations, the significance of green propellants such as hydroxylammonium nitrate, hydrogen peroxide (H2O2) and ammonium dinitramide has become more pronounced. In this paper, recent developments in ceramic microthrusters and the associated ceramic microfabrication techniques are reviewed. The characteristics of green propellants are examined, followed by the evaluation of previous attempts to incorporate green propellants into ceramic microthrusters. This has further unveiled the possibilities of green and clean space missions in the future.

Physical and thermal impact of lead free ballistic modifiers

Preparation of solid propellants based on nitrocellulose was carried out. Modifications on the prepared propellants took place on the basis of environmental aspects. The effect of different ballistic modifiers (bismuth (III) citrate, bismuth sub nitrate and bismuth (III) oxide in comparison with lead oxide) on thermal decomposition behavior of nitrocellulose based propellants was studied. DTA and DSC thermal analysis were conducted to evaluate the performance of the modified propellants. Results from DTA and DSC were used to calculate the apparent activation energy. The relation between the obtained activation energy and the extent of conversion was investigated using DSC data treated with Friedman model. Thermal conductivity, bulk density and specific gravity were also measured to study the effect of burning rate modifier on physical properties of propellant. Results showed that the conventional propellant modifier lead oxide could be replaced by bismuth salts while maintaining the main thermal and physical properties of propellant without a significant shift in its thermal behavior.

Micro-fibre based Porous Composite Propellants with High Regression Rates

Harnessing energy at micro-scale from high energy sources has gained significance in recent times for space propulsion and other applications. Conventional solid rocket propellants have advantages in terms of being efficient, compact and safe to handle, though with much lower regression rates as compared to solid explosives. An approach to high regression rates in composite propellants is demonstrated in the present work by the enhancement of fuel-oxidiser interaction, and by the incorporation of micro-scale porosity into the propellant grain. The porous polystyrene-ammonium perchlorate grain designed in this work, based on electrospun micro-fibres and aqueous impregnation, exhibits burning rates more than 25 times as compared to the non-porous grain. Such high regression rates using insensitive propellant compositions have practical implications in the development of micro-thrusters, and in gas generating devices such as MAV launch systems and turbine starters. Detailed preparatory procedure, characterisation techniques, and flame regression studies are included in this paper.

Directed energy missions for planetary defense

Directed energy for planetary defense is now a viable option and is superior in many ways to other proposed technologies, being able to defend the Earth against all known threats. This paper presents basic ideas behind a directed energy planetary defense system that utilizes laser ablation of an asteroid to impart a deflecting force on the target. A conceptual philosophy called DE-STAR, which stands for Directed Energy System for Targeting of Asteroids and exploration, is an orbiting stand-off system, which has been described in other papers. This paper describes a smaller, stand-on system known as DE-STARLITE as a reduced-scale version of DE-STAR. Both share the same basic heritage of a directed energy array that heats the surface of the target to the point of high surface vapor pressure that causes significant mass ejection thus forming an ejection plume of material from the target that acts as a rocket to deflect the object. This is generally classified as laser ablation. DE-STARLITE uses conventional propellant for launch to LEO and then ion engines to propel the spacecraft from LEO to the near-Earth asteroid (NEA). During laser ablation, the asteroid itself provides the propellant source material; thus a very modest spacecraft can deflect an asteroid much larger than would be possible with a system of similar mission mass using ion beam deflection (IBD) or a gravity tractor. DE-STARLITE is capable of deflecting an Apophis-class (325m diameter) asteroid with a 1- to 15-year targeting time (laser on time) depending on the system design. The mission fits within the rough mission parameters of the Asteroid Redirect Mission (ARM) program in terms of mass and size. DE-STARLITE also has much greater capability for planetary defense than current proposals and is readily scalable to match the threat. It can deflect all known threats with sufficient warning.

Plasma Ignition Response for LOVA Gun Propellant at Low Loading Densities

The ignition of LOVA gun propellants at low loading densities is of interest for future gun systems with insensitive modular charges. The ignition system in conventional propellant systems is traditionally optimized for high loading densities. There is a need to be able to control the ignition for the envelop of the system, to reduce unburnt propellant at low loading densities and, at the same time, avoid pressure waves at high loading densities. Plasma ignition is interesting for this application, with the possibility to vary ignition energy and with the plasma plume penetrating deep into the propellant bed to achieve uniform ignition. This paper presents results from plasma ignition experiments in a 45-mm gun system, using NL008, an RDX-based CAB-LOVA propellant. The plasma igniter is a capillary jet type, driven by a pulse-forming network. A zero-dimensional model for the plasma igniter has been developed and will be discussed in the context of the experiments.

Giant and Tunable Mechanical Impulse of Energetic Nanocrystalline Porous Silicon

In this paper, a nanocrystalline porous silicon-based propulsion system is designed, and its explosion impulse is tuned by varying the propulsion design parameters and etching duration. The porous silicon is ignited by an electrical current passed through 100-nm-thick aluminum film deposited on the unpolished side of the wafer and the ignited porous silicon caused strong explosion, which destroys the chip into tiny fragments. The explosion impulse in the system can reach about at optimal conditions, which is two orders stronger than the impulse produced by conventional propellants (ZakarE., “Technology Challenges in Solid Energetic Materials for Micro Propulsion Applications” U.S. Army Research Lab. Rept. ARL-TR-5035, Nov. 2009). It is also shown that, by varying the etching time, which is an important factor that determines the porous layer thickness and the volume of nanocrystallite, the strength of the impulse can be tuned. Furthermore, a linear increasing trend of the explosion impulse with etching time is observed, which can be explained as the results of heat trapping, materials confinement, and the increasing number of reaction centers.

Utilization of Residual Helium to Extend Satellite Lifetimes and Mitigate Space Debris

A new electric propulsion device concept takes advantage of residual helium gas that is trapped in the chemical propellant feed system and currently unused at end of life. The helium ion thruster provides additional propellant resources to extend satellite lifetimes and transfer geostationary orbit space assets to ultrasafe disposal orbits. The predicted capability, if fully allotted to the disposal, allows for perigee heights above the geostationary altitude that are one order of magnitude greater than existing international guidelines of 250 km. Furthermore, the proposed helium ion thruster concept makes the classic propellant gauge uncertainty problem moot, as satellite operators could use all of their conventional propellant for nominal station-keeping operations. The helium ion thruster concept therefore mitigates future space debris arising from depleted assets in the geostationary orbit belt through both aggressive orbit raising and depressurization of satellites at end of life. An analysis of the helium ion thruster theoretical performance shows that the device could raise the altitude of an end-of-life 2500 kg, 5 kW spacecraft by 2200 km in two months using 2 kg of residual helium.

Mechanical Properties of Smokeless Composite Rocket Propellants Based on Prilled Ammonium Dinitramide

AbstractAmmonium dinitramide (ADN) is a high performance solid oxidizer of interest for use in high impulse and smokeless composite rocket propellant formulations. While rocket propellants based on ADN may be both efficient, clean burning, and environmentally benign, ADN suffers from several notable disadvantages such as pronounced hygroscopicity, significant impact and friction sensitivity, moderate thermal instability, and numerous compatibility issues. Prilled ADN is now a commercially available and convenient product that addresses some of these disadvantages by lowering the specific surface area and thereby improving handling, processing, and stability. In this work, we report the preparation, friction and impact sensitivity and mechanical properties of several smokeless propellant formulations based on prilled ADN and isocyanate cured and plasticized glycidyl azide polymer (GAP) or polycaprolactone‐polyether. We found such propellants to have very poor mechanical properties in unmodified form and to display somewhat unreliable curing. However, by incorporation of octogen (HMX) and a neutral polymeric bonding agent (NPBA), the mechanical properties of such smokeless formulations were significantly improved. Impact and friction sensitivities of these propellants compare satisfactorily with conventional propellants based on ammonium perchlorate (AP) and inert binder systems.

Experiments on the Combustion Characteristics of Deterrent-Coated Propellants and Their Application in Traveling Charge Propulsion

The combustion characteristics of deterrent-coated propellants and their application in a 105 mm caliber gun are investigated experimentally. The mixture of deterrent-coated and conventional single base propellant particles are fired in a 100 ml closed bomb with loading densities of 0.1 g/cm3 and 0.2 g/cm3, respectively. The measured pressure time profiles indicate that the deterrent can retard the combustion of coated propellants, and the delays are related to the pressure and the thickness of the deterrent. The optimized deterrent-coated propellants are attached to the tailfin of a 105 mm caliber armor piercing projectile (APP) as traveling charge, and fired in a 105 mm caliber gun. The firing results indicate that the traveling charge of deterrent-coated propellants can improve the muzzle velocity of projectiles by 2.8% and 4.2% under maximum pressure of 330 Mpa and 400 Mpa, respectively. The pressure time curves measured in the chamber indicate that pressure plateau effects can be expected by the combustion delay of deterrent-coated propellants.

Performance Characteristics of a DME Propellant Arcjet Thruster

This paper describes the influence of cathode configuration on performance of an arcjet thruster using dimethyl ether (DME) propellant. DME, an ether compound, has suitable characteristics for a space propulsion system; DME is storable in a liquid state without being kept under a high pressure, and requires no sophisticated temperature management such as a cryogenic device. DME can be gasified and liquefied simply by adjusting temperature whereas hydrazine, a conventional propellant, requires an iridium-based particulate catalyst for its gasification. In this study, thrust of a 1-kW class DME arcjet thruster is measured at a discharge current of 13 A, DME mass flow rates ranging 15 to 60 mg/s under three cathode configurations: flat-tip rods of 2 and 4 mm in diam. and 4-mm-diam. rod having a cavity of 2 mm in diameter. Thrust measurements show that thrust is increased with propellant mass flow rate. Among the tested cathodes, the flat-tip rod of 4 mm in diam. with 55 mg/s DME flow rate yielded the highest performance: specific impulse of 330 s, thrust of 0.18 N, discharge power of 1400 W and specific power of 25 MJ/kg.