Characterization of electric solid propellant pulsed microthrusters

Electric solid propellants are an attractive option for space propulsion because they are ignited by applied electric power only. In this work, the behavior of pulsed microthruster devices utilizing such a material is investigated. These devices are similar in function and operation to the pulsed plasma thruster, which typically uses Teflon as propellant. A Faraday probe, Langmuir triple probe, residual gas analyzer, pendulum thrust stand and high speed camera are utilized as diagnostic devices. These thrusters are made in batches, of which a few devices were tested experimentally in vacuum environments. Results indicate a plume electron temperature of about 1.7 eV, with an electron density between 1011 and 1014 cm-3. According to thermal equilibrium and adiabatic expansion calculations, these relatively hot electrons are mixed with ~2000 K neutral and ion species, forming a non-equilibrium gas. From time-of-flight analysis, this gas mixture plume has an effective velocity of 1500-1650 m/s on centerline. The ablated mass of this plume is 215 µg on average, of which an estimated 0.3% is ionized species while 45±11% is ablated at negligible relative speed. This late-time ablation occurs on a time scale three times that of the 0.5 ms pulse discharge, and does not contribute to the measured 0.21 mN-s impulse per pulse. Similar values have previously been measured in pulsed plasma thrusters. These observations indicate the electric solid propellant material in this configuration behaves similar to Teflon in an electrothermal pulsed plasma thruster --Abstract, page iv

Characterisation and thrust measurements from electrolytic decomposition of Ammonium Dinitramide (ADN) based liquid monopropellant FLP-103 in MEMS thrusters

Although Ammonium Dinitramide (ADN) has been targeted as a potential green monopropellant in future space vehicles, its application potential in Micro-electrical-Mechanical System (MEMS) thrusters or microthrusters have seldom reported in open literature. In this paper, electrolytic decomposition of Ammonium dinitramide (ADN)-based liquid monopropellant FLP-103 were carried out in open chamber and MEMS thrusters fabricated from Poly-Dimethylsiloxane (PDMS) to characterize the power consumption. Two thrust measurement methods were employed to investigate the electrolytic decomposition of FLP-103 in MEMS microthrusters. The results show that the monopropellant can be successfully ignited at room temperature through 80V,0.1A (8W) using copper wire as electrodes. In the current thruster design, low thrust was obtained at FLP-103 flowrate of 40µl/min but it generated the highest specific impulse, Isp, among all the flowrates tested. The experiments successfully demonstrated the potential application of electrolytic decomposition of FLP-103 in MEMS thrusters

Performance Measurements of Electric Solid Propellant in an Ablative Pulsed Electric Thruster

Electric solid propellants are advanced solid chemical rocket propellants that can be controlled (ignited, throttled and extinguished) through the application and removal of an electric current. These propellants may also be used for electric in-space propulsion, specifically in the ablative pulsed plasma thruster. In this paper, we will investigate the performance of an electric solid propellant operating in an ablation-fed pulsed plasma device by use of an inverted pendulum micro-Newton thrust stand. Namely, the impulse-per-pulse and the specific impulse of the device using the electric solid propellant will be reported for test runs of 100 pulses and energy levels of 5, 10, 15 and 20 J. Further, the device will also be tested using the current state-of-the-art pulsed plasma thruster propellant, polytetrafluoroethylene. The performance of each propellant will be compared for each energy level using an identical setup and apparatus. This comparison of performance between propellants in a controlled setting will allow for better understanding of previous experimental observations

On-a-chip microdischarge thruster arrays inspired by photonic device technology for plasma television

This study shows that the practical scaling of a hollow cathode thruster device to MEMS level should be possible albeit with significant divergence from traditional design. The main divergence is the need to operate at discharge pressures between 1-3bar to maintain emitter diameter pressure products of similar values to conventional hollow cathode devices. Without operating at these pressures emitter cavity dimensions become prohibitively large for maintenance of the hollow cathode effect and without which discharge voltage would be in the hundreds of volts as with conventional microdischarge devices. In addition this requires sufficiently constrictive orifice diameters in the 10µm – 50µm range for single cathodes or <5µm larger arrays. Operation at this pressure results in very small Debye lengths (4 -5.2pm) and leads to large reductions in effective work function (0.3 – 0.43eV) via the Schottky effect. Consequently, simple work function lowering compounds such as lanthanum hexaboride (LaB6) can be used to reduce operating temperature without the significant manufacturing complexity of producing porous impregnated thermionic emitters as with macro scale hollow cathodes, while still operating <1200°C at the emitter surface. The literature shows that LaB6 can be deposited using a variety of standard microfabrication techniques

Impulse Measurements of Electric Solid Propellant in an Electrothermal Ablation-Fed Pulsed Plasma Thruster

Electric solid propellants are advanced solid chemical rocket propellants that can be controlled (ignited, throttled and extinguished) through the application and removal of an electric current. These propellants are also being considered for use in the ablative pulsed plasma thruster. In this paper, the performance of an electric solid propellant operating in an electrothermal ablation-fed pulsed plasma thruster was investigated using an inverted pendulum micro-Newton thrust stand. The impulse bit and specific impulse of the device using the electric solid propellant were measured for short-duration test runs of 100 pulses and longer-duration runs to end-of-life, at energy levels of 5, 10, 15 and 20 J. Also, the device was operated using the current state-of-the-art ablation-fed pulsed plasma thruster propellant, polytetrafluoroethylene or PTFE. Impulse bit measurements for PTFE indicate 10020 N-s at an initial energy level of 5 J, which increases linearly by ~30 N-s/J with increased initial energy. Measurements of the impulse bit for the electric solid propellant are on average lower than PTFE by 10% or less. Specific impulse for when operating on PTFE is calculated to be about 450 s compared to 225 s for the electric solid propellant. The 50% reduction in specific impulse is due to increased mass ablated during operation with the electric solid propellant relative to PTFE

Microthruster Experimental Analysis

The purpose of this study was to design an experimental apparatus that could accurately test and measure the thrust efficiency of conical shaped microthrusters with varying divergence half angles. The experimental apparatus measured the thrust of micronozzles for various feed pressures in vacuum (to duplicate in space operation) as well as ambient. Calibration experiments confirmed the force measurement accuracy of the setup while gas thrust experiments were used to help determine the most efficient divergence half angle. Thrust results were compared to findings from two separate scientific studies that sought to optimize microthruster nozzles using CFD software

Comparison of numerical predictions of the supersonic expansion inside micronozzles of micro-resistojets

none3The present work provides a numerical investigation of the supersonic flow inside a planar micronozzle configuration under different gas rarefaction conditions. Two different propellants have been considered, namely water vapor and nitrogen, which relate to their use in VLMs (the former) and cold gas microthrusters (the latter), respectively. Furthermore, two different numerical approaches have been used due to the different gas rarefaction regime, i.e. the typical continuum Navier–Stokes with partial slip assumption at walls and the particle–based Direct Simulation Monte Carlo (DSMC) technique. As a result, under high–pressure operating conditions, both water and nitrogen flows supersonically expanded into the micronozzle without chocking in combination with a linear growth of the boundary layer on walls. However, when low–pressure operating condition are imposed and a molecular regime is established inside the micronozzle, a very rapid expansion occurred close to the nozzle exit in combination with a strong chocking of the flow and a micronozzle quality reduction of about 40%. Furthermore, water exhibited specific higher specific impulse than nitrogen above 60%.openDe Giorgi M. G., Fontanarosa D., Ficarella AntonioDe Giorgi, M. G.; Fontanarosa, D.; Ficarella, Antoni

Chemical Microthruster Options

Chemical propulsion systems with potential application to microsatellites are classified by propellant phase, i.e. gas, liquid, or solid. Four promising concepts are selected based on performance, weight, size, cost, and reliability. The selected concepts, in varying stages of development, are advanced monopropellants, tridyne(TM), electrolysis, and solid gas generator propulsion. Tridyne(TM) and electrolysis propulsion are compared vs. existing cold gas and monopropellant systems for selected microsatellite missions. Electrolysis is shown to provide a significant weight advantage over monopropellant propulsion for an orbit transfer and plane change mission. Tridyne(TM) is shown to provide a significant advantage over cold gas thrusters for orbit trimming and spacecraft separation

Experimental investigation of a 2.5 centimeter diameter Kaufman microthruster

A 2.5-centimeter-diameter Kaufman electron bombardment microthruster was fabricated and tested. The microthruster design was based on the 15-centimeter-diameter SERT 2 and 5-centimeter-diameter Lewis experimental thruster designs. The microthruster with a two-grid system, operating at a net accelerating potential of 600 volts and an accelerator potential of 500 volts, produced a calculated 445 micronewton thrust when it was run with a 9-milliampere beam current. A glass grid was initially used in testing. Later a two-grid system was successfully incorporated. Both the propellant utilization efficiency and the total power efficiency were lower than for large-size advanced thrusters, as expected; but they were sufficiently high that 2.5-centimeter thrusters show promise for future space applications. Total power of the microthruster with an assumed 7-watt hollow-cathode neutralizer was less than 30 watts at a thrust level of 445 micronewton (100 Nu LBf). The hollow cathode was operated at zero tip heater power for power requirement tests

Development of solid propellant microthrusters

Ph.DDOCTOR OF PHILOSOPH