Monopropellant Research Papers - Academia.edu (original) (raw)

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- Chemistry, Analytical Chemistry, Kinetics, Decomposition
- by Mohammad Fatehi
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- Propulsion, Monopropellant, Interval arithmetic, Uncertainty analysis

Scaling and throttling of combustion devices are important capabilities to demonstrate in development of liquid rocket engines for NASA's Space Exploration Mission. Scaling provides the ability to design new injectors and injection elements with predictable performance on the basis of test experience with existing injectors and elements, and could be a key aspect of future development programs. Throttling is the reduction of thrust with

- by Marlow Moser
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- Engineering, Aerospace Engineering, Materials Science, Mechanics

The burning rate of a propellant is one of the most desired pieces of information for rocket motor design. Propellant burning rate is known to be linked to the microscale flame structures located just above the propellant surface. Flame structure and burning rate for an ammonium perchlorate composite propellant depend in large part on three factors: ammonium perchlorate particle size, propellant formulation, and pressure. Propellant burning rates are in general higher with decreasing AP particle size and increasing pressure. When the microscale flame structures sit higher, on average, above the propellant surface, the propellant will have slower burning rates due (in part) to decreased heat feedback to the propellant surface. The addition of burning rate modifiers to the propellant will also change the flame structure, and therefore the burning rate. Currently, propellants are developed using iterations of mixing and testing to obtain burning rates and physical parameters for computer models. However, this method is not optimal due to the large amount of time and cost involved with this highly empirical approach. Ideally, modelers would be able to make a priori predictions of formulation burning rates, but we are far from that currently. Modelers do desire to create high-fidelity computer models to simulate burning rocket propellants, and much progress has been made in recent years; however, relatively little is known about the actual flame structure in composite propellants which has had limited advances. Knowledge of the variation of flame structure with pressure and propellant formulation will not only assist in the validation of these high-fidelity computer models but will also provide insight to propellant formulators as they seek to use alternate ingredients and methods. This chapter seeks to describe the current data we have on the flame structures in ammonium perchlorate composite propellants and how microscale flame structure affects global burning rate. We review the status of current

- by Steven Son
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- Materials Science, Microscale Chemistry, Monopropellant, Propellant

- by Steven Son
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- Materials Science, Microscale Chemistry, Monopropellant, Propellant

In solid rocket propellants, nano-sized catalysts are expected to be more effective than their micron-sized counterparts due to their higher surface area and increased contact with the oxidizer. However, propellant processing becomes more difficult and ultimate mechanical properties can be negatively impacted as catalyst size is reduced. One proposed solution to these issues is to encapsulate a nano-sized catalyst inside the oxidizer. We have previously created composite particles with nano-sized iron oxide catalyst encapsulated in fine ammonium perchlorate (AP). In this paper we explore the effect of the composite particles on the burning rate and flame structure of an AP-based composite propellant. The propellant containing the composite nano-iron oxide/AP particles was compared to a baseline propellant without catalyst, a propellant formulated with micron-sized catalyst, and a propellant with nano-sized catalyst mixed directly. All catalyzed propellants contained the same catalyst mass fraction. High-speed (5 kHz) OH planar laser-induced fluorescence (PLIF) and high-speed surface imaging were used to examine differences in flame structure and coarse crystal burning characteristics and determine how the encapsulated catalyst influences the global burning rate. It was found that the burning rate of the propellant with the encapsulated catalyst was 90% higher than that of the baseline propellant, 44% higher than the propellant with micron-scale catalyst, and 15% higher than that of the propellant with nano-sized catalyst added directly. Our results indicate that the high global burning rate of the propellant with the encapsulated catalyst is due to an accelerating effect on the fine AP/binder matrix burning rate, assumed to be caused by the intimate contact between the fine AP and catalyst. At elevated pressures (4.4-7.1 atm) flame structure and burning surface morphology change as the catalyst size and location is changed. The coarse crystals are observed to protrude more from the burning surface as the catalyst size decreases. The microscale flame structure is observed to transition from jet-like to lifted arches as the pressure and catalyst size increase.

- by Steven Son
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- Mechanical Engineering, Chemical Engineering, Materials Science, Catalysis
- by Robert Masse
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- Aerospace Engineering, Computer Science, Propulsion, Monopropellant

The NASA Space Technology mission Directorate's (STMD) Green Propellant Infusion Mission (GPIM) Technology Demonstration Mission (TDM) will demonstrate an operational AF-M315E green propellant propulsion system. Aerojet-Rocketdyne is responsible for the development of the propulsion system payload. This paper statuses the propulsion system module development, including thruster design, system design and system component materials compatibility testing. Major system components of the propulsion system module include: propellant tank, latch valve, service valve and thruster valve. All system components, except the thruster valve, are flight proven (TRL 9) for hydrazine propellant; Status is given on modifications of these components to ensure that all internal wetted surfaces are compatible with the AF-M315E propellant. The culmination of this program will be high-performance, green AF-M315E propulsion system technology at TRL 7+, with components demonstrated to TRL 9, ready for direct infusion to a wide range of applications for the space user community. Nomenclature EM = Engineering model ESPA = EELV secondary payload adapter GPIM = Green Propellant Infusion Mission HAN = Hydroxyl ammonium nitrate I sp

- by Robert Masse
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- Engineering, Aerospace Engineering, Systems Engineering, Propulsion

Cold gas propulsion systems offer today's CubeSats a relatively simple, minimal V propulsion solution. CubeSat mission needs have been identified that would be enhanced or enabled if significant V were available in conjunction with preserving the control authority offered by cold gas propulsion systems. Examples include large scale orbit transfer for constellation deployment, proximity operations, formation-flying, de-orbit, orbit maintenance, attitude control, and momentum management. To meet these needs, Aerojet is offering the MRS-142, a 1U blow-down CubeSat High-impulse Adaptable Monopropellant Propulsion System (CHAMPS) that delivers a more than five-fold increase in total impulse compared to similarly-packaged cold gas systems. A four-thruster array, providing threeaxis attitude control as well as single-axis V, is integrated into a monolithic piston propellant tank doubling as the primary structure. To satisfy varied mission-specific propulsion system requirements, the CHAMPS design supports adjustment of the thrust level and minimum impulse via easily changeable fluid resistors and system operating pressure. Thrust vector orientation is tailored through simple modifications of the thruster nozzles. The final flight system is designed to be fully compliant with Air Force Space Command Manual 91-710.

- by Robert Masse
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- Engineering, Aerospace Engineering, Astrobiology, Propulsion
- by Michael Houts
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- Engineering, Propulsion, Nuclear power

Current liquid propellants are either cryogenic (storage constraints), either extremely toxic (handling problem). For this purpose, alternative storable propellant are currently being investigated in this work. The sprays generated by a like-impingement configuration have to be ignited. Because the couple of propellants chosen is non hypergolic, the energy of ignition was brought by a torch igniter. The injection conditions of both the propellants and the torch proved to be compatible to ignite the reactants quickly and stabilize the combustion downstream of the impingement point.

- by bastien boust
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- Chemistry, Combustion, Monopropellant, Propellant

The numerical procedure for the burning of Ammonium Perchlorate (AP) with a Fuel-Binder (Hydroxyl Terminated Polybutadience HTPB) heterogeneous propellant is presented. This model accounts for the one-step reaction mechanism for the primary diffusion flame between the decomposition products of the Binder (B) and the oxidizer AP and allowed for the complete coupling between the gas-phase physics, the condensed-phase physics, and the unsteady non-uniform regression of the propellant surface. The parameters used in this model are fitted to experimental data for the combustion of AP/Binder. The propagation of the unsteady non-planer regression surface is described, using the Essentially-Non-Oscillatory (ENO) scheme with the aid of the level set strategy. The Alternating-Direction-Implicit (ADI) solver is employed to solve the full Navier-Stokes equations in the gas phase. The results show the effect of various parameters on the surface propagation speed, flame structure, and the burning surface geometry. A comparison between the computational and experimental results is presented.

- by mohamed hegab
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- Materials Science, Combustion, Microscale Chemistry, Solver

The utilization of liquid oxygen/liquid methane couple (LOX/LCH4) as a potential candidate to substitute hypergolic propellants and hydrazine in the next future propulsion systems has arisen an increasing interest due to the advantages offered in terms of low environmental impact, re-usability, cooling capabilities and relatively high specific impulse [1]. In this perspective, the Italian Aerospace Research Center manages the "HYPROB" research program, cofunded by the Italian Research and University Ministry, that has the objective to improve the national capabilities into developing engines, fed by hydrocarbons, that could be successfully applied as propulsion units for third stages of launchers for space exploration. The "HYPROB" program led to the realization of a LOX/LCH4 engine named "DEMO-0A", a 30 kN thrust class demonstrator, technologically representative of a regenerative thrust chamber assembly of an expander engine [2]. The present paper describes the results of the numerical simulations performed by means of the EcosimPro software, aimed at reproducing the operative conditions, both cold flow both firing, of the regenerative thrust chamber "DEMO-0A". An assessment of the capabilities of the software in predicting the behaviour of the demonstrator by modelling it with a 1-D approach and by considering different wall heat exchange semiempirical correlations has been done by comparing numerical results and the available experimental data gathered during both cold flow both firing test campaigns.

Monopropellant thruster is one of the most propulsion system types developed in the space industry. This system uses a single type of propellant that reacts in porous medium catalytic packed bed to generate thrust in the form of hot gases. The last decade, green propellant hydrogen peroxide (H2O2), also known as High Test Peroxide (HTP), thanks to its low cost and easy to store as liquid, is used as an alternative solution of hydrazine which is very toxic and not environmentally friendly. In the current study, hydrogen peroxide monopropellant thruster is investigated for application in the future satellites. A numerical simulation is performed using the Computational Fluid Dynamics (CFD) software ANSYS Fluent in order to simulate fluid flow of hydrogen peroxide in thruster, and the finite volume method was employed for resolving the governing equation. Species transport model is applied in the single-phase reaction simulation using the Eddy Dissipation model (EDM) for turbulence-che...

- by Othman Bendermel
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- Monopropellant, Hydrogen Peroxide, Propellant

As a system becomes more complex, the uncertainty in the operating conditions increases. In such a system, implementing a precise failure analysis in early design stage is vital. However, there is a lack of applicable methodology that shows how to implement failure analysis in the early design phase to achieve a robust design. The main purpose of this paper is to present a framework to design a complex engineered system resistant against various factors that may cause failures, when design process is in the conceptual phase and information about detailed system and component is unavailable. Within this framework, we generate a population of feasible designs from a seed functional model, and simulate and classified failure scenarios. We also develop a design selection function to compare robust score for candidate designs, and produce a preference ranking. We implement the proposed method on the design of an aerospace monopropellant propulsion system.

- by Kai Goebel
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- Engineering, Computer Science, Propulsion, Failure Analysis

The article presents alternative metal-supported catalysts for decomposition of the highest-class hydrogen peroxide: 98%+ (Type 98 HP, according to MIL-PRF-16005F). The aim of this study was the experimental investigation of an alternative solution for decomposition of 98%+ hydrogen peroxide, strictly for chemical propulsion. High-porosity open cell metal foams have been identified as structures with great potential. Low density, good mechanical and thermal properties, availability of various materials and alloys as well as new technologies of manufacturing, make metal foam a potential solution for many different propellants, not only hydrogen peroxide. Open cell NiCrAl foam has been processed to prepare several catalysts, with different content and dispersion of the active phase. Cleaning and drying were performed to prepare carriers for further processing: wet impregnation, slow drying, and calcination. Simple drop tests with 98% hydrogen peroxide have been conducted to estimate a...

- by Paweł Surmacz
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- Materials Science, Catalysis, Aerospace, Monopropellant

The paper summarizes the Space Technology Department at the Institute of Aviation goals for small spacecraft propulsion technology and link them to areas of propulsion research in Poland in general at the nearest future. A brief review of currently being investigated issue in the newly opened laboratory is presented as current and near term research, focused specifically on: concentrated solutions of H2O2 (stabilized and HTP - High Test Peroxide class) preparation, its long term storage possibilities, catalytic decomposition, “drop tests” for fuel ignition investigation, its use as “green” monopropellant or in hybrid rocket motors as the oxidant. By employing catalytic decomposition of HTP, auto-ignition of the solid fuel in a hybrid rocket is possible. When HTP is properly decomposed, the decomposition products alone release enough heat and the gases have suitable temperature (even over 700°C, depending on the concentration of the solution of peroxide) to provide efficient thrust. ...

- by Paweł Surmacz
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- Materials Science, Hydrogen Peroxide

A concept was evaluated of using nitrous oxide as (1) a monopropellant in thrusters for space suits and spacecraft and (2) a source of breathable gas inside space suits and spacecraft, both by exploiting the controlled decomposition of N2O into N2 and O2. Relative to one prior monopropellant hydrazine, N2O is much less toxic, yet offers comparable performance. N2O can be stored safely as a liquid at room temperature and unlike another prior monopropellant hydrogen peroxide does not decompose spontaneously. A prototype N2O-based thruster has been demonstrated. It has also been proposed to harness N2O-based thrusters for generating electric power and to use the N2 + O2 decomposition product as a breathable gas. Because of the high performance, safety, and ease of handling of N2O, it can be expected to be economically attractive to equip future spacecraft and space suits with N2O-based thrusters and breathable-gas systems.

- by Robert Zubrin
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- Engineering, Spacecraft Propulsion, Monopropellant, Nitrous Oxide

A study was initiated to investigate propulsion stage and mission architecture options potentially enabled by fission energy. One initial concept is a versatile Nuclear Thermal Propulsion (NTP) system with a maximum specific impulse of 900 s and a maximum thrust (per engine) of 15 klbf. The system assumes a monopropellant stage (hydrogen), and is designed to also provide 300 lbf of thrust (potentially split between multiple thrusters) at an Isp > 500 s for orbital maneuvering and station keeping. Boost pumps are used to assist with engine decay heat removal and low thrust engine burns, and to compensate for partial tank depressurization during full thrust engine burns. Potential stage assembly orbits that take full advantage of launch vehicle payload mass and volume capabilities are being assessed. The potential for using NTP engines to also generate a small to moderate amount of electrical power is also being evaluated. A first generation versatile NTP stage could enable 8 of 9 upcoming opportunities for short (less than 24 month) round trip human missions to Mars. A second generation versatile NTP is under consideration that could potentially provide a maximum specific impulse of 1800 s at 15 klbf, and enable ambitious missions throughout the solar system. The second generation NTP system under consideration would also allow a choice of volatiles to be used as propellant. This would potentially allow in-situ resources such as water, ammonia, methane, or other compounds to be used directly as propellant by the second generation engine.

- by Michael Houts
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- Aerospace Engineering, Propulsion, Monopropellant, Thrust
- by Michael Houts
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- Aerospace Engineering, Materials Science, Propulsion, Monopropellant

- by Vinayak MelaRam Malhotra
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- Hydrogen, Combustion, Monopropellant, Optimal Specific Impulse

The Monopropellant Hydrazine Propulsion system is one of the most widely used types of single-agent propulsion systems to control the position or correction of satellites in orbits. This system consists of combustion chamber subsystems (catalyst bed, catalyst, nozzle, and cap), fuel and fuel tank, high-pressure tank, control valves, and interface pipes. In this paper, the MPHP system (as a case study) is described in detail, and then critical risks are identified by creating FMECA tables on the case study in the design phase. Based on the proposed FMCEA flowchart, potential failure modes are identified. In the next step, decisions and corrective actions are formulated regarding the inherent failures of the system. Finally, the necessary measures to reduce the risks will be taken according to the system's failure modes, and the reduction of the identified risks to an acceptable level is presented. The attained results show that the catalyst decomposition chamber, catalyst bed, inlet flow control valve, and propellant management facilities units have the highest risk index values (RPN), respectively. For this purpose, corrective measures have been suggested for each of these.

P e y x i d~B a s e d Propulsion ~i p t t d l s Department-' V M S a w NaaoMlLsbaptaierr A l~, N e w M e d c o 87185sndUvetm0n,CDPfomia 94550 p 6~k) 8 r i h E a n w~, k r~k s~t e s~e p r r m b n t a c~~s W Nuclear Securib, AdrnlntslraUon undw Conbad DE-ACOCggA(bS000. Q Sandia National laboratories Issued by SandiaNational Laboratories, v t e d for the United States DepQmnt of Enagy by Sandia Corporation. NOTICE: Tlus report was prepad as an account of work sponsored by an agency ofthe United States Government Neither the Umted States Government, nor any agency thereof, nor any of then employees, nor any of thnr contractors, subconir8c&n, or their mployca make any wmanty, express or nnpkcd, or mume any legal lrability or r q o m b~l r t y for the accuracy, wmpletemss, or usehtlness of any information, appamtus, product, or praeess dtsclosed, or represent that its use would aot inhnge privately owned rights. Reference henin to any qmciiic commercial product, pmcess, or savice by trade name, mkmwk, mawlfkmer, or otherwise, doas not necesgarily conshtutc or nnply endorsement, reaannnendation, or fswring by h e United States Gov-t , any agency thereof, or any of their mh.adors w submntmcton. The views and opinions e x p r d herem do not necmsdy stnte or reflect tbose of the Umted States Government, any agency thereof, or my of their contractors. Pnnted in the United States of America This report bas bem reproduced directly fmm tho best available copy.

- by Brian Melof
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- Engineering, Propulsion, Monopropellant, Hydrogen Peroxide
- by Brandie Rhodes
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- Astrobiology, Propulsion, Monopropellant, Satellite

In monopropellant system, hydrogen peroxide is used with catalyst to create an exothermic reaction. Catalyst made of silver among the popular choice for this application. Since the catalyst used is in porous state, the effect of its porosity in the hydrogen peroxide monopropellant thruster performances is yet unknown. The porosity changes depending on factors including catalyst pact compaction pressure, bed dimension, and type of catalyst used. As researches on this topic is relatively small, the optimum porosity value is usually left out. The performance of the thruster indicated by the pressure drop across the catalyst bed. Porosity of the catalyst bed adds additional momentum sink to the momentum equation that contributes to the pressure gradient which lead to pressure loss inside thruster. The effect of porosity influences the performance and precision of the thruster. Study of the pressure drop by the catalyst bed requires a lengthy period and expensive experiments, however, nu...

- by NORAZILA BINTI OTHMAN, Dr.
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- Materials Science, Catalysis, Porosity, Monopropellant
- by Greg Mungas
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- Engineering, Spacecraft Propulsion, Monopropellant, Nitrous Oxide
- by Greg Mungas
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- Engineering, Spacecraft Propulsion, Monopropellant, Nitrous Oxide

- by Norazila Othman
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- Materials Science, Catalysis, Porosity, Monopropellant

This paper aims at giving a analytical approach of the stability analysis of control pressure components in pneumatic domain. Some authors have given some results in linear case especially in the hydraulic context [Meritt 1960, Mac Cloy et al 1980, Margolis 1997, Alirand et al. 2002]. The structure and the technologies of commonly used pneumatic pressure control component are firstly introduced, and then a non-linear model is proposed as the basis for the stability analysis. The classical linear method, based on the tangent linearized model is applied to determine the analytical stability conditions using Routh-Hurwitz criteria. The root loci are finally studied according to a set of equilibrium points which corresponds to various conditions of use. In this last part, an existing pressure regulator is used as an example and allows the development of an analysis concerning the dynamic performances on realistic basis. To conclude, the proposed approach gives a set of design rules which enables the system parameters to be sized according to the required dynamic performances and stability.

- by Xavier Brun
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- Design, Modeling, Stability Analysis, Root Locus

A liquid propellant alumina microthruster with an integrated heater, catalytic bed and two temperature sensors has been developed and tested using 30 wt. % hydrogen peroxide. The temperature sensors and the catalytic bed were screen-printed using platinum paste on tapes of alumina that was stacked and laminated before sintering. In order to increase the surface of the catalytic bed, the platinum paste was mixed with a sacrificial paste that disappeared during sintering, leaving behind a porous and rough layer. Complete evaporation and combustion, resulting in only gas coming from the outlet, was achieved with powers above 3.7 W for a propellant flow of 50 µl/min. At this power, the catalytic bed reached a maximum temperature of 147°C. The component was successfully operated up to a temperature of 307°C, where it cracked.

- by Greger Thornell
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- Engineering, Physics, Materials Science, Sintering

A combustion assembly capable of continuously burning monopropellant and bipropellant liquid fuels at pressures up to 80 bars (1145 psig) was designed and constructed. The assembly is based on a liquid propellant strand burner where a manifold maintains small positive differential pressures on the fuel to maintain a steady supply into the reaction vessel. Optical ports enable direct visualization of the flame and will allow for future spectroscopic and imaging studies of the flame. The strand burner design was tested using nitromethane with both air and inert environments in the reaction vessel. Continuous combustion was sustained for almost 8 min in air (34 bars/500 psig) and more than 6 min in N2 (70 bars/1000 psig). A unique outcome from the initial testing of this device is the ability to ignite liquid nitromethane in an inert environment without the use of a pilot flame started in air.

- by Michael Klassen
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- Engineering, Materials Science, Combustion, Medicine
- by Michael Houts
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- Aerospace Engineering, Environmental Science, Astrobiology, Propulsion

The fundamental capability of Nuclear Thermal Propulsion (NTP) is game changing for space exploration. A first generation NTP system could provide high thrust at a specific impulse (Isp) above 900 s, roughly double that of state of the art chemical engines. Characteristics of fission and NTP indicate that useful first generation systems will provide a foundation for future systems with extremely high performance. The role of a first generation NTP in the development of advanced nuclear propulsion systems could be analogous to the role of the DC-3 in the development of advanced aviation systems. Nomenclature CFEET = Compact Fuel Element Environmental Test DOE = Department of Energy HAT = NASA Human Architecture Team HIP = Hot Isostatic Press

- by Michael Houts
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- Engineering, Aerospace Engineering, Propulsion, Optimal Specific Impulse

In the current scenario of space propulsion, liquid propellants have significantly proved useful in the upper stage rocket engines. Over the past couple decades, the world had inclined positively towards cryogenic fuel(s) viz., liquid oxygen and liquid hydrogen due to their high specific impulse. A higher specific impulse implies lower duration to achieve design cruise velocity for a given rocket initial and instantaneous mass. Liquid hydrogen and liquid oxygen as fuel and oxidizer can generate one of the highest enthalpy release in combustion, producing a specific impulse of up to 450 seconds at an effective exhaust velocity of 4.4 kilometres per second. Whereas, selected disadvantages are encountered in the form of storage and production. This indicates overdependence on cryogenic propellants and has necessitated the active research effort for better alternatives. As an interesting alternative, the combination of Dinitrogen Tetroxide (N 2 O 4) and Monomethyl Hydrazine (MMH) have been used for many space applications owing to an extreme storage stability and hypergolic nature. Present study aims to express the effect of hydrogen-based compounds on the rocket performance. Four distinctive compounds from two groups of hydrogen-based compounds are tested with the varying oxidizer and fuel proportions to obtain a new, cost-effective and user-friendly composition that can be prepared at room temperature. The investigation attempt and explains the effect of hydrogen based energetic propellants using N 2 O 4 and MMH as the base composition for upper stage performance. The work is motivated by the need of efficient space operations with attractive propulsive alternatives to minimize overdependence on cryogenics, which will ultimately result in cost effectiveness. Various energetic materials were tested with the base composition by using standard NASA-CEA complex chemical equilibrium model. The performance was evaluated in terms of variation in specific impulse and characteristic velocity both of which are significant parameters. To, validate the practical utility, the role of chamber pressure, supersonic area ratio and optimal Oxidizer to fuel ratio (O/F) was determined. The work led to two interesting findings, a composition of beryllium hydride with base composition for high performance of rockets and the negative impact of hydrogen on liquid propellants.

- by TRIPARNA RAY
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- Hydrogen, Combustion, Monopropellant, Optimal Specific Impulse

- by TRIPARNA RAY
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- Hydrogen, Combustion, Monopropellant, Optimal Specific Impulse

This paper describes a direct-injection configuration of a monopropellant-powered actuator that is intended to provide high-energy-density actuation for a self-powered positionor force-controlled human-scale robot. The proposed actuator is pressurized by a pair of solenoid injection valves (each of which control the flow of a monopropellant through a catalyst pack and directly into the respective side of a pneumatic-type cylinder), and depressurized via a three-way hot-gas proportional exhaust valve. A controller is described that coordinates the control of the two solenoid propellant injection valves, together with the control of the proportional hot-gas exhaust valve, in order to provide actuator force tracking. Experimental results are presented that validate the effectiveness of the force-control approach. Finally, energetic performance of the proposed actuator is experimentally assessed and shown to provide an energetic figure of merit, an order of magnitude greater than that of a battery-powered servomotor approach.

Hydrazine (N2H4) is produced at industrial scale from the partial oxidation of ammonia or urea. The hydrogen content (12.5 wt%) and price of hydrazine make it a good source of hydrogen fuel, which is also easily transportable in the hydrate form, thus enabling the production of H2 in situ. N2H4 is currently used as a monopropellant thruster to control and adjust the orbits and altitudes of spacecrafts and satellites; with similar procedures applicable in new carbon-free technologies for power generators, e.g. proton-exchange membrane fuel cells. The N2H4 decomposition is usually catalysed by the expensive Ir/Al2O3 material, but a more aff ordable catalyst is needed to scale-up the process whilst retaining reaction control. Using a complementary range of computational tools, including newly developed micro-kinetic simulations, we have derived and analysed the N2H4 decomposition mechanism on the Cu(111) surface, where the energetic terms of all states have been corrected by entropic terms. The simulated temperature-programmed reactions have shown how the pre-adsorbed N2H4 coverage and heating rate aff ect the evolution of products, including NH3, N2 and H2. The batch reactor simulations have revealed that for the scenario of an ideal Cu terrace, a slow but constant production of H2 occurs, 5.4% at a temperature of 350 K, while the discharged NH3 can be recycled into N2H4. These results show that Cu(111) is not suitable for hydrogen production from hydrazine. However, real catalysts are multi-faceted and present defects, where previous work has shown a more favourable N2H4 decomposition mechanism, and, perhaps, the decomposition of NH3 improves the production of hydrogen. As such, further investigation is needed to develop a general picture.

- by Alberto Roldan
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- Chemical Engineering, Chemistry, Catalysis, Decomposition

This paper presents a dynamic model of the interior ballistics of an experimental liquid propellant powered rifle. The liquid propellant powered rifle described utilizes a mixture of hydroxyl ammonium nitrate and hydrocarbon fuel to replace gunpowder typically used in such firearms. The motivation for such a development is to discard the need for a shell casing whereby carrying only propellant and bullets will reduce both the mass and volume per shot carried by the soldier. A first-principles dynamic model of the interior ballistics is derived as a compressible fluid power problem with the chemical liberation of heat within the chamber modeled via a condensed-phase reaction rate law. The model is used to predict the overall performance in terms of ballistic kinetic energy as well as draw design insight regarding the role of friction, chamber geometry, and the profile of chamber pressure with respect to time. Simulation results are presented as well as preliminary experimental result...

- by Eric Barth
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- Engineering, Mechanical Engineering, Materials Science, Design

Research and development of green oxidizers and green fuels as a possible replacement for ammonium perchlorate (NH 4 ClO 4 , AP) and hydrazine (N 2 H 4) respectively has been increased considerably in the recent years. AP and hydrazine are the oxidizer and fuel entities, and used in solid and liquid rocket motors respectively. AP is highly toxic and led to adverse health effects, while hydrazine is carcinogenic in nature. AP is in use from the last several decades for rocket and space shuttle propulsion, while hydrazine is used in upper stage liquid propelled rocket motors. It's a tough task to replace AP with the currently available green oxidizers; since their ballistic properties are weaker when compared to AP and also they can't be successfully deployed in a solid rocket motor at present. Some important available solid green oxidizers are ammonium nitrate (AN), ammonium dinitramide (ADN), hydroxyl ammonium nitrate (HAN), and hydrazinium nitroformate (HNF). However, AN is one of the cheap and readily available oxidizer, and has great potential to use in solid/ liquid rocket motors. Tremendous progress has been envisaged till now, and more progress will be there in the coming future over the development of AN based green energetic materials (GEM's). A concise overview has been presented over the development of phase stabilized ammonium nitrate (PSAN) and AN/KDN based green oxidizers in the present review paper.

- by PRATIM KUMAR
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- Chemistry, Defence Technology

- by Greger Thornell
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- Engineering, Physics, Materials Science, Sintering

Nomenclature = interfacial area for heat transfer, 1 = gas specific heat, 60 J/mole K = solid specific heat, 1 .0 J/g K = heat transfer coefficient, 210 J/m 2-s K = effective thermal conductivity of catalyst bed, 0.30 T s J/m-s K = gas pressure, atm = gas constant = rate of surface reaction per unit surface area = amount of catalyst surface area per unit volume of reactor, 1.2 xlO 5 m 2 /£ = gas phase temperature, K = solid temperature, K = gas phase velocity, m/s = mole fraction of monopropellant in gas = axial distance along the catalyst bed, cm = void fraction of catalyst bed, 0.5 = heat of reaction, 80 k J/mole = gas density = solid density, 4 kg/f

Composite propellants feature a di¨usion §ame. The size of oxidizer particles leverage some combustion properties (mainly, burning rate and pressure sensitivity) along with §ame structure. Macroscopic combustion features are strictly related to those events occurring inside the gas phase and close to the burning surface. The §ame of nonaluminized composite energetic materials is considered and a simpli¦ed combustion model is tested for this case. Combustion of a laminate propellant with varying lamina size is simulated. The benchmark consists of some movies taken from ammonium perchlorate (AP) / hydroxyl-terminated polybutadiene (HTPB) propellant combustion with a high-speed video camera. Three di¨erent powder sizes are used in propellant manufacturing.

- by Luciano Galfetti
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- Materials Science, Combustion, Composite Material, Monopropellant

The combustion properties of propellants like ethane / nitrous oxide mixtures that have the potential to substitute hydrazine or hydrazine / dinitrogen tetroxide in chemical propulsion systems are investigated. In support of CFDsimulations of new rocket engines powered by green propellants ignition delay times of ethane / nitrous oxide mixtures diluted with nitrogen have been measured behind reflected shock waves at atmospheric and elevated pressures, at stoichiometric and fuel-rich conditions aimed for the validation of reaction mechanism. In addition, ignition delay time measurements of ethene / nitrous oxide mixtures and ethane / O 2 / N 2-mixtures with an O 2 :N 2 ratio of 1:2 as oxidant at the same level of dilution are shown for comparison. Finally, the ignition delay time predictions of a recently published reaction mechanism by Glarborg et al. are compared with the experimental results.

- by Uwe Riedel
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- Chemistry, Monopropellant, Nitrous Oxide, Propellant

Проведен анализ используемых ранее и современных систем наддува баков окислителя и горючего ракет-носителей, двигательные установки которых используют жидкий кислород и керосин. Исследованы возможности использования жидкого аммиака в системах наддува топливных баков в качестве рабочего тела. Рассмотрены характеристики аммиака и тепловые аспекты его разложения на водород и азот. Предложены схемы, использующие тепло факела двигателя и тепло твердотопливного газогенератора. Показана высокая эффективность аммиачных СН на примере I ступени РН «Зенит».

- by Богдана Крысько
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- Materials Science, Monopropellant, Kerosene, Liquid Oxygen

This paper describes the design and energetic characterization of an actuator designed to provide enhanced system energy and power density for self-powered robots. The proposed actuator is similar to a typical compressible gas fluid-powered actuator, but pressurizes the respective cylinder chambers via a pair of proportional injector valves, which control the flow of a liquid monopropellant through a pair of catalyst packs and into the respective sides of the double-acting cylinder. This paper describes the design of the proportional injection valves and describes the structure of a force controller for the actuator. Finally, an energetic characterization of the actuator shows improvement relative to prior configurations and marked improvement relative to stateof-the-art batteries and motors.

- by Kevin Fite
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- Engineering, Mechanical Engineering, Manufacturing Engineering, Monopropellant
- by Claude Joyner
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- Aerospace Engineering, Materials Science, Propulsion, Monopropellant

The propellant synthesis community is constantly looking for green alternative monopropellants. Energetic ionic liquids have several attractive properties such as high energy content, high bulk density, low vapor pressure, high thermal stability, wide liquidus range, low corrosiveness, low toxicity, and ease of handling. The combustion characteristics of an energetic ionic liquid hydroxyethylhydrazinium nitrate (HEHN) was conducted in a pressurized chamber. The performance of HEHN was compared to that of the monopropellant Otto fuel II (OF-II) typically used for torpedo-propulsion. A liquid strand combustion study was performed in an atmosphere of air and nitrogen with chamber pressures varying from 10 to 90 bar. Regressing surface profiles and subsequent burning rates were obtained at different chamber pressures. A B-type thermocouple of 46 μm wire diameter was used to measure the monopropellant flame temperature of HEHN. Thermogravimetric analysis was performed to study the thermal decomposition of HEHN to understand its thermal stability and Fourier transform infrared spectroscopy (FTIR) was utilized to determine the possible reasons behind the high burning rates of HEHN. The gains in the specific impulse and density specific impulse coupled with enhanced burning rates and reasonable thermal stability are expected to establish HEHN as a frontrunner for propulsion and power-generation in oxygen-deficient scenarios.

- by Umakant Swami
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- Mechanical Engineering, Chemical Engineering, Materials Science, Ionic Liquid

This paper presents a liquid-fuel powered pneumatic actuator appropriate for human-scale autonomous robotics. The motivation for this work is the development of a lightweight actuation system with system energy and power densities significantly greater than a DC motor and battery combination. For the scale of interest, the design tradeoffs between complexity of an energy conversion system and fuel specific energy density make many conventional approaches inappropriate. Conventional actuation, such as a battery powered DC motor system, does not possess adequate energy storage to perform significant amounts of mechanical work for significant periods of time autonomously. A system design comparison for a six-degree of freedom walking robot is presented which compares the expected performance of a monopropellant powered actuation system to a battery powered DC motor system. Results from this comparison demonstrate the viability of such an approach and indicate significant increases in energy and power densities. A single-degree of freedom manipulator was constructed to demonstrate this approach. Experimental results corroborate the expected energetic advantages.

- by Eric Barth
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- Engineering, Robotics, Computer Science, Control Engineering