Academic literature on the topic 'Hydrazine propellants'

A review of the literature pertaining to hypergolic fuel systems, particularly using hydrazine or its derivatives and hydrogen peroxide, has been conducted. It has been shown that a large effort has been made towards minimizing the risks involved with the use of a toxic propellant such as the hydrazine. Substitution of hydrazines for nontoxic propellant formulations such as the use of high purity hydrogen peroxide with various types of fuels is one of the major areas of study for future hypergolic propellants. A series of criteria for future hypergolic propellants has been recommended, including low toxicity, wide temperature range applicability, short ignition delay, high specific impulse or density specific impulse, and storability at room temperature.

Abstract This paper describes the recent theoretical and experimental research by the Netherlands Organisation for Applied Scientific Research (TNO) into green replacements for hydrazine, hydrazine derivatives and nitrogen tetroxide, as propellants for in-space propulsion. The goal of the study was to identify propellants that are capable of outperforming the current propellants for space propulsion and are significantly less hazardous for humans and the environment. Two types of propellants were investigated, being monopropellants and bipropellants. The first section of the paper discusses the propellant selection. Nitromethane was found to be the most promising monopropellant. As bipropellant, a combination of hydrogen peroxide (HP) and ethanol was selected, where the ethanol is rendered hypergolic with hydrogen peroxide. The second part of the paper describes the experimental verification of these propellants by means of engine testing. Initiation of the decomposition of nitromethane was found to be problematic, hypergolic ignition of the hydrogen peroxide and ethanol bipropellant however was successfully demonstrated.

A wide range of propellants, and propulsion systems in space exploration by aircrafts or space vehicles was studied, developed, investigated, and commercialized. Liquid, solid, or hybrid propellants have been used for rocket’s launches. In this review, a consistent definition of space propulsion systems, including solid, liquid and hybrid has been given with up-to-date state of developments. A comparison of their performances was made by theoretical and experimental specific impulses. On the other hand, ammonium perchlorate and hydrazine were used as propellants for rocket’s launches and for satellite’s maneuverings; respectively. However, their high toxicity and their storage problem pushed researchers and scientists to investigate and develop other eco-friendly, propellant systems, so called “green propellants”, for launch or reaction control systems of satellites.

The vacuum specific impulse, density vacuum specific impulse, and solid exhaust products were examined for several propellant formulations based on the pyrophoric material triethylaluminum (TEA) using CEA thermodynamics code. Evaluation of TEA neat and mixed with hydrocarbon fuels with LOX, N2O, N2O4, liquefied air, and HNO3were performed at stoichiometry. The vacuum specific impulse of neat TEA with N2O is comparable to that of nitric acid with the same, but the N2O formulation will produce slightly less solid products during combustion. Additionally, N2O-TEA propellants have vacuum specific impulses and density vacuum specific impulses within 92.9% and 86.7% of traditional hydrazine propellant formulations under stoichiometric conditions.

Water can be utilized as spacecraft propellant to dramatically reduce the environmental impact of constructing and operating a satellite. In this work, a multi-mode chemical-electrical propulsion system, in which water was used as the propellant in both high thrust chemical and high specific impulse electrical maneuvres, was studied. This type of system allows the spacecraft architecture community to divest from traditional propellants such as hydrazine and xenon, thus reducing the production of highly toxic chemicals and dramatically reducing the carbon footprint of propulsion systems. Water has the lowest toxicity, carbon footprint, and price of any current or proposed propellant, and has been shown in laboratory testing to be a feasible alternative compared to traditionally used propellants. The unique role it can play across multiple spacecraft subsystems suggests that the commercial adoption of water as a propellant will reduce cost and mass while also reducing the environmental impact of the satellites of tomorrow. This technology has the ability to enable the development of modular, multifunctional, competitive, and environmentally friendly spacecraft architectures.

The aim of this work is to analyze the state of the art in the development and use of pollution-free (“green”) propellants in low-thrust jet engines used as actuators of spacecraft stabilization and flight control systems and to adapt computational methods to the determination of “green”-propellant engine thrust characteristics. The monopropellant that is now widely used in the above-mentioned engines is hydrazine, whose decomposition produces a jet thrust due to the gaseous reaction products flowing out of a supersonic nozzle. Because of the high toxicity of hydrazine and the complex technology of hydrazine filling, it is important to search for its less toxic substitutes that would compare well with it in energy and mass characteristics. A promising line of this substitution is the use of ion liquids classed with “green” ones. The main components of these propellants are a water solution of an ion liquid and a fuel component. The exothermic thermocatalytic decomposition of a “green” propellant is combined with the combustion of its fuel component and increases the combustion chamber pressure due to the formation of gaseous products, which produces an engine thrust. It is well known that a “green” propellant itself and the products of its decomposition and combustion are far less toxic that hydrazine and the products of its decomposition, The paper presents data on foreign developments of “green” propellants of different types, which are under test in ground (bench) conditions and on a number of spacecraft. The key parameter that governs the efficiency of the jet propulsion system thrust characteristics is the performance of the decomposition and combustion products, which depends on their temperature and chemical composition. The use of equilibrium high-temperature process calculation methods for this purpose is too idealized and calls for experimental verification. Besides, a substantial contribution to the end effect is made by the design features of propellant feed and flow through a fine-dispersed catalyst layer aimed at maximizing the monopropellant-catalyst contact area. As a result, in addition to the computational determination of the thrust characteristics of a propulsion system under design, its experimental tryout is mandatory. The literature gives information on the performance data of “green”-propellant propulsion systems for single engines. However, in spacecraft control engine systems their number may amount to 8–16; in addition, they operate in different regimes and may differ in thrust/throttling characteristics, which leads to unstable propellant feed to operating engines. To predict these processes, the paper suggests a mathematical model developed at the Institute of Technical Mechanics of the National Academy of Sciences of Ukraine and the State Space Agency of Ukraine and adapted to “green”-propellant engine systems. The model serves to calculate the operation of low-thrust jet engine systems and describes the propellant flow in propellant feed lines, propellant valves, and combustion chambers. To implement the model, use was made of the results of experimental studies on a prototype “green”-propellant engine developed at Yuzhnoye State Design Office. The analysis of the experimental results made it possible to refine the performance parameters of the monopropellant employed and obtain computational data that may be used in analyzing the operation of a single engine or an engine system on this propellant type in ground and flight conditions

GSLV is a “geosynchronous satellite launch vehicle” which aims at launching space objects into GTO- Geosynchronous transfer orbit. This paper provides a brief idea of the chemistry of propellants used in all three stages and strap-on motors of geosynchronous satellite launch vehicles and their combustion mechanism. It includes all series of GSLVs launched till now and a description of criteria for choosing fuel. This paper also includes a brief description on hydrazine which is a basic part of so many propellant combinations. Furthermore, it includes basic reasons which result in unsuccessful ignition especially at cryogenic. It also emphasizes on reasons for modification in fuels over time and advancement in efficiency obtained due to respective modified fuels, further, giving an overview of future ideas in the development of propellants.

A thruster, designated LE111, utilizes hydrazine and MON3 as fuel and oxidizer and produces a thrust of 111 N at nominal operating conditions with an Isp of up to 328 s. The thruster is throttleable through the use of separate fuel and oxidizer metering valves and is able to transition to a hydrazine monopropellant mode at any point during its operation. Completed qualification testing has demonstrated a versatile operating box that includes chamber pressure-mixture ratio excursions, heated propellants up to 341 K, and gaseous helium ingestion through propellant flow paths. Overall, the thruster has achieved 15,455 s of accumulated on-time and a maximum continuous burn time of 6,000 s. The thruster achieves its performance through a novel micro-coaxial gas-gas injection scheme and an innovative regeneratively cooled combustion chamber, which uses the oxidizer as the working fluid. The thruster's capabilities and performance recorded during the completed qualification testing are described, and its design is outlined. This paper is published with the permission of the authors granted to 3AF – Association Aéronautique et Astronautique de France (www.3AF.fr) organizer of the Space Propulsion International conference.

Abstract A number of propellant/thruster combinations are under development in recent years that aim to replace the prevailing hydrazine-driven reaction control thrusters with less hazardous substances (“green propellants”). With some of these systems already in orbit, characterizing their contamination potential in a space environment becomes relevant. In this paper we discuss experiments on plume induced contamination from a novel propene/nitrous oxide bipropellant thruster, including high-speed imaging, SEM-EDS analysis, QCM measurements and in-situ mass spectrometry. Additional measurements of combustion chamber pressure complement the overall characterization of the thruster performance in a high vacuum environment. The main findings of this exploratory study are strong indications of solid and liquid phase particles being ejected from the nozzle, which will be investigated in a subsequent phase of the activity.

Thesis (Ph. D.)--Chemistry and Biochemistry, Georgia Institute of Technology, 2007.
Committee Chair: Prof. Charles L. Liotta; Committee Member: Prof. Arthur J. Ragauskas; Committee Member: Prof. Charles A. Eckert; Committee Member: Prof. John D. Muzzy; Committee Member: Prof. Rigiberto Hernandez.

Monopropellants are widely used for propulsion and gas generation applications. The current state of art monopropellant systems are based on hydrazine propellants. While hydrazine has a strong heritage as a versatile monopropellant, some of the inherent problems associated with hydrazine like toxicity, high vapor pressure and associated storage and handling cost have been a major concern. These concerns led to the exploration of nontoxic and better performing alternate propellants, among which Hydroxylammonium Nitrate (HAN) is a promising candidate and scores over other monopropellants in terms of insensitivity, toxicity and volatility. Some of the other projected advantages of HAN formulations over hydrazine include its lower crystallization point, higher density and volumetric impulse. The current practice of decomposing HAN is based on available technology used for hydrazine decomposition. However main challenges for HAN thrusters in contrast to hydrazine are material constraints imposed by the higher temperature levels experienced in HAN thrusters and need of high temperature tolerant catalysts. The catalysts under consideration should be both temperature tolerant and resistant to any possible poisoning from the decomposition products of HAN. There have been few reports of late on the development of suitable catalysts for HAN. Besides, unlike hydrazine, HAN propellants are multi-component, undergoes multistage decomposition adding complexity to the decomposition mechanism. Current work pertains to the development of a high temperature tolerant active catalyst for HAN decomposition. The issues related to catalyst design, catalyst characterization, reactivity to HAN based monopropellants, performance estimation and its decomposition kinetics are extensively pursued in this work. Special tests were developed to measure catalyst’s capability to sustain extreme chemical and thermal conditions witnessed in HAN thruster. A novel cobalt doped cerium oxide (CeCo) catalyst was prepared via co-precipitation route followed by slip casting for pelletization. Besides, as an improvement over this design, iridium coated bifunctional catalyst was prepared by wet impregnation of CeCo catalyst pellets. To optimize composition of cobalt doped ceria catalyst, the catalytic activity of different compositions of catalyst was checked and a simultaneous investigation into the possible cause of their reactivity was conducted. The cobalt doped ceria catalyst was compared for reactivity and sustainability against conventional Ir/γ-Al2O3 catalyst. Thermoanalytical methods were employed to determine onset of decomposition temperature, rate of decomposition and exothermicity of the reaction. An inhouse designed batch reactor dedicated to examining durability and sustainability of the catalyst was used. To investigate any change caused by exposure to extreme exothermic decomposition of monopropellant, catalyst samples were subjected to physical and chemical characterization before and after decomposition. The characterization techniques used for the catalyst were Scanning Electron Microscopy/Energy Dispersive Spectroscopy, X-ray Diffraction Spectroscopy and X-ray Photoemission Spectroscopy. The activity of cobalt doped ceria was evident in thermal analysis since it evinced a low temperature and a high exothermic decomposition with HAN. Further, the new catalyst retained its physical integrity and was extremely resistant to catalyst poisoning during HAN decomposition in batch reactor while iridium-based alumina supported catalyst couldn’t sustain its activity either due to poisoning or severe attrition. A specific composition of cobalt was optimized in ceria for optimum performance in terms of lowering of decomposition temperature and endurance in relation to iridium catalyst. The structural failure of Ir/γ-Al2O3 pellets and physical integrity of ceria catalyst during batch reactor studies for multiple trials were demonstrated through SEM studies. The activity of new catalyst in decomposing HAN was found to be a function of Ce3+ presence in ceria matrix as determined from XPS. Thermal poisoning due to high decomposition temperature experienced in HAN thruster was simulated by exposing the catalysts in a high temperature furnace and modifications if any was followed using thermal analysis techniques. The exceptional ability of iridium metal in initiating decomposition of HAN at a significantly low temperature, leads us to incorporate iridium as a reactive metal over cobalt doped cerium oxide catalyst to form eventually a bifunctional catalyst. This bifunctional catalyst contained both Lewis acid and base sites to facilitate an optimised redox decomposition of HAN. Thermal analysis and batch reactor studies showed promising features for this catalyst. XPS studies carried out on these bifunctional catalysts show predominantly surface presence of active metal compared to bulk of catalyst. Iridium in bifunctional catalyst showed signs of oxidation during HAN decomposition along with an overall decrease in quantity suggesting bleeding of oxidized iridium due to high decomposition temperatures. However, Ce3+ quantities in bifunctional catalyst followed the pattern similar to the one observed in cobalt doped ceria catalyst for HAN decomposition. Despite deactivation of iridium due to oxidation, the promotional effect of iridium in bifunctional catalyst was indisputably evident in these studies. Since HAN has a 33% positive oxygen balance it is used in combination with a suitable fuel for higher performance. The stoichiometrically balanced ternary systems obviously produce higher temperatures and extreme conditions to which the catalysts have to be tested. The ternary system tried out in the present study composed to HAN, methanol and water in the ratio 70:15:15 (w/w). The performance analysis of the catalyst was restricted to batch reactor since thermal studies with slow heating rates does not allow a simultaneous decomposition due to the differential boiling points of the components. However, the batch reactor was preheated to facilitate minimum time lag between ternary HAN system injection and its subsequent decomposition. The uncatalyzed ternary HAN decomposition showed lower energy outputs in comparison to binary systems. However, the performance of catalysed decomposition of ternary HAN system was remarkably superior to HAN binary system. Initially, the performance of both the catalysts was promising and comparable. However, Ir/γ-Al2O3 catalyst got disintegrated and deactivated much earlier as demonstrated by XRD and XPS. While the cobalt doped ceria catalyst sustained its activity and provided a consistent performance along at least 50 injections of HAN ternary system. The response of cobalt doped ceria catalyst to HAN ternary system was more or less similar to the binary system. XRD and XPS studies after the decomposition showed retention of Ce3+, the active component of the ceria-based catalyst, even after multiple trials which also point to its resistance to deactivation even in harsh reaction conditions produced by actual propellant system comprising of fuel, HAN and water. Chemical kinetics provides an important insight into the decomposition mechanism and helps in the performance analysis of propellant compositions. The present work explores the chemical kinetics involved in the thermal and catalytic decomposition of HAN using established methods in thermal analysis. The kinetic parameters were determined from thermo-gravimetric analysis by using isoconversional methods. Two isothermal methods namely, Popescu-Ortega’s method (integral method) and Friedman’s method (differential method) were used for kinetics estimation. The integral method failed due to its limitations in multi-step reactions when a physical step of water evaporation always preceded HAN decomposition in slow heating thermal analysis studies. Whereas, differential method worked well after smoothing the TGA data. The values obtained from differential method were complemented with the results obtained from EGA (evolved gas analysis) for which a DTA-TG-FTIR hyphenated system was used. Primary product species identified for HAN decomposition were N2O, HNO3 and NO2 for all cases. The reliability of isoconversional method adopted here was initially verified for thermal decomposition for which kinetics parameters are available. The variable activation energy calculated as a function of conversion was used to describe the variation in decomposition mechanisms for both the catalysts tested in the present study. The variation in the concentration of species with change of catalyst was prominent and further demonstrated the superiority of ceria catalyst developed in this work. The EGA profiles further substantiated the proposed decomposition mechanism and suggested reasons for robustness of ceria catalysts over iridium catalyst. The new class of cerium oxide catalyst developed during this work looks highly promising thanks to its unique features like absence of attrition, resistance to poisoning, high temperature tolerance, propensity to initiate decomposition reaction at low temperature with high exothermicity and preference towards thermodynamic products leading to a low molecular mass product stream.

AbstractThe substitution of the toxic hydrazine in current high-altitude rocket engines like upper stages or reaction control thrusters by green propellants is a major key driver in the current technology development of rocket propulsion systems. Operating these kind of rocket engines at high-altitude leads to a sudden pressure drop in the liquid propellants during their injection into the combustion chamber with a near-vacuum atmosphere prior to ignition. The resulting superheated thermodynamic state of the liquid causes a fast and eruptive evaporation which is called flash boiling. The degree of atomisation is important for a successful ignition and a secure operation of the rocket engine. The development and operation of a cryogenic high-altitude test bench at DLR Lampoldshausen enables the systematical experimental characterization of cryogenic flash boiling due to its ability to adjust and control the injection parameters like temperature, pressure or geometry. Several test campaigns with liquid nitrogen (LN2) were performed using two optical diagnostic methods: First, flash boiling LN2 spray patterns were visualised by means of high-speed shadowgraphy and, secondly, we determined the droplet size and velocity distributions in strongly superheated LN2 sprays with the help of a laser-based Phase Doppler system (PDA). The experimental data generated within these measurement campaigns provide defined boundary conditions as well as a broad data base for the numerical modelling of cryogenic flash boiling like e.g. the publications [8, 9].

Abstract An experimental investigation has been carried out for small liquid bi-propellant thrusters of 490 & 1500 N levels. These thrusters have to operate for more than 100 sec in continuous and pulse mode. In this case, film cooling is the primary mode of thruster cooling. The thruster uses hydrazine based propellant as fuel and N2O4 as oxidiser. Film cooling is carried out by injecting a fraction of fuel from an injector periphery. Unlike impinging type injection elements are used for core flow. The thruster’s shell used for testing was made of stainless steel and di-silicide coated C103 material. A 1D heat transfer model was developed for predicting the thruster outer wall temperature. The experimental investigation was carried for different film cooling percentage and injector configuration was modified for each case. Thermocouples were mounted on top & bottom side of shell for temperature measurement. Infrared camera also used for recording temperature in the test. Based on experimental investigation, effective film cooling percentage for optimum thruster performance has been estimated for two thrust levels and these studies also helped in validating the heat transfer model.

During the start-up of the propulsion system of a satellite or spacecraft, the opening of the tank isolation valve will cause the propellant to flow into an evacuated feedline and slam against a closed thruster valve. This filling process, called priming, can cause severe pressure peaks that could lead to structural failure. In the case of monopropellants such as hydrazine, also, the risk of adiabatic compression detonation must be taken into account in the design of the feedline subsystem. The phenomenon of priming involves complex two-phase flow: the liquid entering the evacuated pipe undergoes flash evaporation creating a vapor cushion in front of the liquid that mixes with the residual inert gas in the line. Moreover, the dissolved pressurizing gas in the liquid will desorb making the priming process difficult to model. In order to study this phenomenon, a new test-bench has been built at DLR Lampoldshausen which allows fluid transient experiments in the same conditions as the operating space system. Tests are performed with water and ethanol at different conditions (tank pressure, vacuum level, pressurizing gas helium vs. nitrogen, etc.). The effect of the geometry is also investigated, comparing different test-elements such as straight, tees, and elbow pipes. The pressure profile is found to be dependent on the geometry and on the downstream conditions. The acoustic wave reflection caused by the pipe geometry and fluid dynamic effects such as the aforementioned desorption and flash evaporation induce a complex pressure profile of the first pressure peak. Finally, numerical simulations of the priming process are performed by means of EcosimPro software in conjunction with European Space Propulsion System Simulation (ESPSS) libraries and results are compared with experiments.