June 16, 2021

Download Ebook Free Electrospray Thrusters For Spacecraft Propulsion

Micromachined Electrospray Thrusters for Spacecraft Propulsion

Micromachined Electrospray Thrusters for Spacecraft Propulsion
Author : Renato Krpoun
Publisher : Unknown
Release Date : 2009
Category :
Total pages :133
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Microfabricated Electrospray Thrusters for a Modular Spacecraft Propulsion System

Microfabricated Electrospray Thrusters for a Modular Spacecraft Propulsion System
Author : Simon Dandavino,Herbert Shea (Physiker)
Publisher : Unknown
Release Date : 2014
Category :
Total pages :230
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Development of Ionic Liquid Multi-mode Spacecraft Micropropulsion Systems

Development of Ionic Liquid Multi-mode Spacecraft Micropropulsion Systems
Author : Steven Paul Berg
Publisher : Unknown
Release Date : 2015
Category :
Total pages :149
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"This dissertation presents work on development of multi-mode specific spacecraft propulsion systems. Specifically, this work attempts to realize a single propellant capable of both chemical monopropellant and electric electrospray rocket propulsion, develop methods to characterize multi-mode propulsion system performance, and realize a system capable of both monopropellant and electrospray propulsion for a small spacecraft. Selection criteria for ionic liquid propellants capable of both monopropellant and electrospray propulsion are developed. These are based on desired physical properties and performance considering use in both propulsive modes. From these insights, a monopropellant mixture of 1-ethyl-3-methylimidazolium ethyl sulfate and hydroxylammonium nitrate is selected and synthesized. Multi-mode spacecraft micropropulsion systems which include a high-thrust chemical mode and high-specific impulse electric mode are assessed. Due to the combination of a common propellant for both propulsive modes, low inert mass, and high electric thrust, the monopropellant/electrospray system has the highest mission capability in terms of delta-V for missions lasting shorter than 150 days. The ionic liquid monopropellant mixture is tested for decomposition on heated platinum, rhenium, and titanium surfaces. It was found that the propellant decomposes at 165 °C on titanium, which is the decomposition temperature of HAN, and 85 °C on platinum. Arrhenius-type reaction rate parameters were calculated from the results and used to develop thruster models. The [Emim] [EtSO4]-HAN propellant mixture is tested in a capillary electrospray emitter and exhibits stable electrospray emission at a nominal extraction voltage of 3400 V. The highest specific impulse attained in these experiments was 412 seconds; however, this could be improved with a more robust feed system design. This data, along with data from the monopropellant decomposition experiment is used to design a multi-mode micropropulsion system using a common propellant and common thruster geometry. This system is capable of ~20-40% greater delta-V capability at a given mission duration compared to a system utilizing separate, state-of-the-art monopropellant and electrospray thrusters"--Abstract, page iv.

Design and Analysis of a Stage-based Electrospray Propulsion System for CubeSats

Design and Analysis of a Stage-based Electrospray Propulsion System for CubeSats
Author : Oliver Jia-Richards
Publisher : Unknown
Release Date : 2019
Category :
Total pages :149
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The standardization of small spacecraft through CubeSats has allowed for more affordable space exploration. This progress in affordability has been limited to Earth orbit due in part to the lack of high [delta]V propulsion systems that are compatible with the small form factor. The ion Electrospray Propulsion System developed at the Space Propulsion Laboratory at the Massachusetts Institute of Technology is a promising technology foundation for a compact, high [delta]V propulsion system. However, the [delta]V output of the propulsion system is limited by the lifetime of individual electrospray thrusters. This thesis presents the design and analysis of a stage-based concept for the ion Electrospray Propulsion System where the propulsion system is composed of a stack of electrospray thruster arrays. The stage-based propulsion system bypasses the lifetime limit of individual electrospray thrusters in order to increase the lifetime of the entire propulsion system. In effect, propulsion capabilities for CubeSats can be advanced without the need for technological developments. With the current performance metrics of the ion Electrospray Propulsion System, deep-space missions with an initial spacecraft form factor of a 3U CubeSat are feasible with current propulsion technology. Mechanisms required for the stage-based system are designed and demonstrated in a vacuum environment. In addition, analytical methodologies for the analysis of stage-based propulsion systems are developed to assist in preliminary mission design as well as provide the framework for autonomous decision making. Finally, applications of a stage-based propulsion system for missions to near-Earth asteroids are explored as well as analytical guidance for the escape trajectory.

Application of Ion Electrospray Propulsion to Lunar and Interplanetary Missions

Application of Ion Electrospray Propulsion to Lunar and Interplanetary Missions
Author : Caleb Wade Whitlock,Massachusetts Institute of Technology. Department of Aeronautics and Astronautics
Publisher : Unknown
Release Date : 2014
Category :
Total pages :123
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High specific impulse electric propulsion systems enable ambitious lunar and interplanetary missions that return a wealth of scientific data. Many of these technologies are difficult to scale down, meaning the spacecraft are relatively massive and expensive. The Space Propulsion Lab (SPL) at the Massachusetts Institute of Technology (MIT) is developing compact, high specific impulse ion electrospray thrusters which do not suffer from the same sizing limitations. The Ion Electrospray Propulsion System (iEPS) is tailored for small spacecraft and can perform high AV maneuvers. This enables a plethora of lunar and interplanetary missions using nanosatellites, which can lead to substantial cost reductions. The main objective of the research presented in this thesis is to develop a guidance and control (GC) architecture for small spacecraft using iEPS modules for main propulsion and attitude control actuation and to evaluate its performance through simulation. The Lunar Impactor mission serves as the primary case study, and the results offer valuable insight into the design of the propulsion system while validating the functionality of the GC algorithm. These methods are extended in a second case study focusing on exploration of a near-earth asteroid.

The Effect of Temperature on the Structure of Colloid Thrusters Beams

The Effect of Temperature on the Structure of Colloid Thrusters Beams
Author : Maria Pau Bravo
Publisher : Unknown
Release Date : 2011
Category :
Total pages :66
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This work investigates the effect of the temperature on the structure of electrospray beams. Electrospraying is a technique used to atomize charged liquids. From the different operational modes, the cone-jet mode is characterized by its stability [1] and is used in several applications as mass spectrometry [2], nanodroplets sputtering [3] or space propulsion (colloid thrusters). Colloid thrusters are electrostatic accelerators of charged particles. The very low thrust that a single needle provides can be used for precision spacecraft positioning and drag cancellation missions. Time-of-flight and retarding potential techniques are used to characterize the ionic liquid 1-Ethyl-3-Methylimidazolium Bis (Trifluoromethylsulfonyl) Imide (EMI-Im) in vacuum, in the range of temperature between 20° C and 50° C. Two populations of particles are studied for several flow rates and temperatures: ions and droplets. The charge, the specific charge, and the retarding potential of the particles are obtained. Time-of flight waves provide time dependent spectra of current I(t) which integration is used to estimate the propulsive parameters (thrust, specific impulse and efficiency) provided by a single emitter-extractor configuration. The thrust and specific impulse delivered by a cone jet depend mainly on the electrical conductivity of the propellant and the flow rate used. By varying these two parameters, a range in thrust between 0.1 [mu] N and 0.6 [mu] N is obtained for an acceleration voltage of 1680 V. Impingement of charged particles on the electrode's surfaces reduces the operational lifetime of the colloid thruster and the total thrust delivered. The retarding potential detector is used to sample current profiles of the electrospray. The spreading of the beam is studied for several flow rates and temperatures of the ionic liquid EMI-Im.

Emiim Wetting Properties & Their Effect on Electrospray Thruster Design

Emiim Wetting Properties & Their Effect on Electrospray Thruster Design
Author : Garrett Donald Reed
Publisher : Unknown
Release Date : 2012
Category : Electric propulsion
Total pages :56
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Recent advances in the development of highly conductive ionic liquids have made them of interest for use as propellant in spacecraft propulsion systems. Electrospray thrusters apply strong electrostatic fields to an ionic liquid in order to extract and accelerate charged particles, producing thrust. The behavior of these ionic liquids as they pass through the components of an electrospray system can have a significant effect on thruster operation. The wetting and adhesion behavior between the ionic liquid propellant and solid materials can be characterized using the surface tension and contact or "wetting" angle. Ideally this angle is a function of the interactions between the solids surface energy, the "free" surface energy or tension of the liquid and the interactions of both with the surrounding medium. Deviation from ideal contact angle behavior can indicate surface inconsistencies, environmental effects or contamination of the solid or liquid. Contact angle and surface tension measurements are presented for the ionic liquid propellant 1-Ethyl-3-methylimidazolium bis(trifluoromethylsulfonyl)imide, called EMIIm or EtMeImTf2N, with respect to various substrate materials and environmental conditions. Analysis of these measurements determines optimum materials and operating conditions for current and future electrospray thruster designs.

Characterization on a Magnetically Levitated Testbed for Electrospray Propulsion Systems

Characterization on a Magnetically Levitated Testbed for Electrospray Propulsion Systems
Author : Fernando Mier Hicks,Massachusetts Institute of Technology. Department of Aeronautics and Astronautics
Publisher : Unknown
Release Date : 2014
Category :
Total pages :82
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Small satellites are changing the space scene dramatically. By drastically reducing costs while still having impressive technological capabilities, their popularity among the space community is increasing at a very fast rate. Propulsion systems for these class of spacecraft are very limited. One promising technology is the ion Electrospray Propulsion System (iEPS) developed at the Space Propulsion Laboratory at MIT. Electrosprays accelerate ions present in the interface between an ionic liquid and vacuum using strong electric fields. Current thrust estimates for the iEPS modules land in the vicinity of tens of [mu]Newtons. Measuring the small thrust produced by the devices is challenging to say the least. This thesis presents the design and development of a Magnetically Levitated Thrust Balance (MLTB) for thrust estimation of the iEPS devices. The MLTB levitates an engineering model of a small satellite using magnetic fields inside a vacuum chamber. The zero friction environment is exploited to measure the minute thrust levels produced by the electrospray thrusters. Additional sensors and actuators that provide added functionality to the instrument are also explained. A fully stand-alone Power Processing Unit (PPU) capable of generating and delivering the high voltage signals needed to drive the thrusters is explained in detail. Test results of charging behavior and lifetime characterization of the emitted current are presented as a preliminary exploration of these processes.

Considerations for a Multi-modal Electrospray Propulsion System

Considerations for a Multi-modal Electrospray Propulsion System
Author : Chase Spenser Coffman,Massachusetts Institute of Technology. Department of Aeronautics and Astronautics
Publisher : Unknown
Release Date : 2012
Category :
Total pages :73
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Micro- and nano-satellites have begun to garner significant interest within the space craft community as economic trends encourage a shift away from larger, stand-alone satellite platforms. In particular, CubeSats have emerged as popular, economic alter natives to traditional satellites which might also facilitate low-cost space access for academia and developing nations. One of the foremost remaining obstacles to the widespread deployment of these spacecraft is the lack of suitable propulsion, which has severely limited the scope of prior CubeSat missions. While these spacecraft have gained traction by virtue of their economical size, the same quality has imposed unique propulsion demands which have continued to elude traditional thruster concepts. The ion Electrospray Propulsion System (iEPS) is a microelectromechanical (MEMS) based electrostatic thruster for space propulsion applications. This technology makes use of ionic liquid ion sources (ILIS) and a porous emitter substrate to obviate the need for cumbersome ancillary components and achieve the spatial and power characteristics that could lend feasibility to active micro/nano-satellite propulsion. This thesis introduces the iEPS concept and highlights the characteristics that make it attractive as a means of CubeSat propulsion. Specifically, its bimodal propulsion characteristics are presented alongside a discussion of the constant power Isp modulation mechanism that makes this unique capability possible. A simple demonstration of the variable Isp concept is reported, and a brief exploration of the performance implications is used to suggest a direction for taking it to operational maturity.

A Fully Microfabricated Two-dimensional Electrospray Array with Applications to Space Propulsion

A Fully Microfabricated Two-dimensional Electrospray Array with Applications to Space Propulsion
Author : Blaise Laurent Patrick Gassend,Massachusetts Institute of Technology. Department of Electrical Engineering and Computer Science
Publisher : Unknown
Release Date : 2007
Category :
Total pages :269
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This thesis presents the design, fabrication and testing of a fully-integrated planar electrospray thruster array, which could lead to more efficient and precise thrusters for space propulsion applications. The same techniques could be used for making arrays to increase throughput in many other electrospray applications. Electrospray thrusters work by electrostatically extracting and accelerating ions or charged droplets from a liquid surface to produce thrust. Emission occurs from sharp emitter tips, which enhance the electric field and constrain the emission location. The electrospray process limits the thrust from a single tip, so that achieving millinewton thrust levels requires an array with tens of thousands of emitters. Silicon batch microfabrication has been used, as it is well suited for making large arrays of emitters. The thruster is made using Deep Reactive Ion Etching (DRIE) and wafer bonding techniques, in a six mask process, and comprises two components. The emitter die with up to 502 emitters in a 113 mm2 area, is formed using DRIE and SF6 etching, and is plasma treated to transport liquid to the tips in a porous black-silicon surface layer. The extractor die incorporates the extractor electrode, a Pyrex layer for insulation, and springs which are used to reversibly assemble the emitter die. This versatile assembly method, with 10 μm RMS alignment accuracy and 1.3 μm RMSD repeatability, allows the extractor die to be reused with multiple emitter dies, and potentially with different emitter concepts than the one presented. The thruster, weighing 5 g, was tested with the ionic liquids EMI-BF4 and EMIIm. Time of flight measurements show that the thruster operates in the ion emission regime most efficient for propulsion, with a specific impulse around 3000 s at a 1 kV extractor voltage. Emission starts as low as 500 V. Currents of 370 nA per emitter have been recorded at 1500 V, for an estimated thrust of 26 nN per emitter or 13 μN total, and a 275 mW power consumption. The thrust efficiency is estimated around 85%. In good operating conditions, the current intercepted on the extractor electrode is well below 1%, increasing to a few percent at the highest current levels. The beam divergence half width half maximum is between 10 and 15°.

CubeSat Attitude Control Using Micronewton Electrospray Thruster Actuation

CubeSat Attitude Control Using Micronewton Electrospray Thruster Actuation
Author : Mark David Van de Loo,Massachusetts Institute of Technology. Department of Aeronautics and Astronautics
Publisher : Unknown
Release Date : 2014
Category :
Total pages :217
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Micronewton electrospray thrusters are a promising new actuator for CubeSat attitude control. Electrospray thrusters have advantages over current state of the art CubeSat attitude actuators in mass, volume, and their ability to produce translational acceleration in addition to control moments. An attitude determination and control system was designed for a 1U CubeSat assuming commercial-off-the-shelf attitude determination hardware components and six electrospray thrusters developed by the MIT Space Propulsion Laboratory. A high fidelity spacecraft dynamics simulation was constructed for analysis of the performance of the ADCS system. Attitude determination was tested with an engineering model of a 1U CubeSat, and the entire ADCS system was tested in simulation. Results of these preliminary tests show the use of electrospray thrusters as attitude actuators to be feasible, although significant work remains to complete a flight-ready ADCS system.

Design and Manufacturing of an Ion Electrospray Propulsion System Package and Passively-fed Propellant Supply

Design and Manufacturing of an Ion Electrospray Propulsion System Package and Passively-fed Propellant Supply
Author : Louis Evan Perna,Massachusetts Institute of Technology. Department of Aeronautics and Astronautics
Publisher : Unknown
Release Date : 2014
Category :
Total pages :130
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Satellites under 500 kilograms have been growing more popular with the miniaturization of high-performance electronics and instruments. Constellations and formations of satellites consisting of thousands of small satellites will enable inexpensive, on-demand, global access to spaceborne assets. The only impediment to the adoption of small satellites and their exploitation in radical new space system architectures is an absence of high-specific-impulse, scalable, benign propulsion options. Available technologies are too resource inefficient for small satellites, too inflexible, or pose a threat to primary launch payloads. An emergent technology, electrospray propulsion, is inherently scalable, benign, applicable to a wide range of mission types, and resource efficient. Research in the MIT Space Propulsion Laboratory over the past decade has been focused on developing robust electrospray propulsion systems scaled to the needs of small spacecraft. The Ion Electrospray Propulsion System (iEPS) is the synthesis of this work and features a fully-integrated power processing unit (PPU), propellant supply, and electrostatic ion accelerator designed for use in CubeSats. To meet the objectives of the iEPS project, development was necessary for all three components. The work described here focused on a redesign of the thruster module package and initial design and testing of a compact, passive propellant supply system. A MEMS package was designed, manufactured, and tested. It comprised and contained critical electrospray components in close, precise proximity and maintained electrical isolation between high voltage electrodes. Additionally, the package provided for structural and electrical attachment interfaces for the PPU and propellant supply. Design rationale is presented and iterative improvements described for both the package components and manufacturing processes. A prototype passive propellant supply system was designed and tested. The results of integration and testing for both components are presented with discussion of challenges and potential improvements.

Spacecraft Charging and Attitude Control Characterization of Electrospray Thrusters on a Magnetically Levitated Testbed

Spacecraft Charging and Attitude Control Characterization of Electrospray Thrusters on a Magnetically Levitated Testbed
Author : Fernando Mier Hicks
Publisher : Unknown
Release Date : 2017
Category :
Total pages :189
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Electrospray thrusters are an attractive technology for small satellite propulsion. A thorough study of the spacecraft charging and attitude control performance of electrospray thrusters acting on a small satellite is presented. The experimental portion of the study employs a new type of magnetically levitated testbed to measure thrust, characterize spacecraft charging phenomena, and demonstrate precision attitude control maneuvers. The testbed magnetically levitates a mockup-satellite in vacuum conditions. The satellite carries thrusters, batteries, radio, and high voltage thruster electronics. The magnetic levitation and high vacuum provide a zero-friction environment for the thrusters to actuate. The thrusters are placed on the satellite in such a way to produce a net torque when fired. The thrusters are fired and the corresponding rotational movement of the satellite is analyzed to calculate thrust. The uncertainty of the thrust measurement is estimated to be ±0.35 [mu]tN (3[sigma]). Theoretical and experimental methods were developed to investigate the spacecraft charging characteristics anticipated to be observed on spacecraft during the operation of electrospray thrusters. These devices can produce positively and negatively charged ion-beams. An electrical model of this configuration was created to predict the charging properties of electrically isolated systems. This model simulates a bipolar electrospray thruster system. Experiments were conducted on the magnetic levitation testbed. The results from the experimental tests demonstrate that neutralization with heavy ionic species is indeed possible. The thrusters are able to fire in a bipolar configuration for long periods of time inducing bounded spacecraft charging in the range from -400V to +600V when emitting currents of about 20 [mu]A. It was found that the presence of low-energy ions produced by the fragmentation of large clusters and the external space plasma play a significant role in the neutralization characteristics. This electrical model was verified by reproducing the experimental results, thus validating its use to estimate bulk spacecraft charging properties. Electrospray thrusters are capable of producing impulse bits in the 10-6 N-s or lower, depending on the configuration. These characteristics have the potential to allow for long term pointing in the arcsecond range or better with practically no jitter. Experimental work is performed on the magnetic levitation testbed to demonstrate the actuation characteristics under non-optimal control. It is found that subject to a noisy attitude sensor and external perturbations, electrospray thrusters are capable of producing pointing accuracies of 22 arcseconds 3[sigma] error during 10 hours on a platform similar in size and mass to a 1U CubeSat. The implementation of such capabilities could complement or eliminate the need of reaction wheels and magnetorquers, especially in missions beyond low earth orbit, while including propulsive capabilities for additional maneuverability in technology development or scientific missions.

Ionic Liquid Ion Source Emitter Arrays Fabricated on Bulk Porous Substrates for Spacecraft Propulsion

Ionic Liquid Ion Source Emitter Arrays Fabricated on Bulk Porous Substrates for Spacecraft Propulsion
Author : Daniel George Courtney,Massachusetts Institute of Technology. Department of Aeronautics and Astronautics
Publisher : Unknown
Release Date : 2011
Category :
Total pages :332
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Ionic Liquid Ion Sources (ILIS) are a subset of electrospray capable of producing bipolar beams of pure ions from ionic liquids. Ionic liquids are room temperature molten salts, characterized by negligible vapor pressures, relative high conductivities and surface tensions lower than water. Compared with the colloid form of electrospray, renowned for its applications to spectroscopy, ILIS yield highly monoenergetic beams composed entirely of ions. In this respect they are similar to Liquid Metal Ions Sources, but offer the ability to emit both positive and negative ions from a benign propellant that remains in the liquid state over a wide range of temperatures. When applied to spacecraft propulsion these sources are very power efficient and yield high specific impulses. Furthermore. the low flow rates and negligible vapor pressures of ionic liquids allow for passive feeding systems which can remain exposed to the vacuum of space. This configuration would remove the need for pressurized propellant tanks or valves, both of which are difficult to miniaturize for small satellites. However; the thrust produced from each emitter is very low, less than 0.1 [mu]N. As a result, compact arrays of active ILIS have been sought since their discovery. If arrays of modest packing density (~ 5 emitters/mm2) could be achieved, ILIS as thrusters would offer a scalable form of propulsion capable of providing useful thrust levels to small satellites with performance comparable to established, but difficult to miniaturize, plasma based ion engines. This research has sought a technique for creating arrays of ILIS from bulk porous substrates as part of an overall process for microfabricating complete thrusters. The thesis includes a survey of potential fabrication methods considering both suitability for forming arrays of ILIS and the ability to integrate each technique within a thruster packaging process. Electrochemical etching is highly selective and can proceed at rates which are limited by mass transport conditions. In this thesis we show how this etching regime can be exploited to smoothly remove material from the surface of a bulk porous metal substrate without damaging the internal pore structure. Dry film photoresists have been identified as a suitable alternative to spin on techniques for porous materials and have been applied within an electrochemical etching process. A two step process for forming arrays of ILIS has been motivated using numerical simulations of the etching process to predict emitter profiles and investigate the impacts of non-uniform etching conditions. These concepts have been applied experimentally using a custom built, automated, etching station capable of repeatedly producing arrays of 480 emitters spaced 500 pm apart on a 1 x 1 cm porous nickel substrate pre-mounted, and aligned, within a silicon thruster package. The emitters are typically 165 [mu]m tall with rounded tips suitable for operation as ILIS. Pulsed voltage conditions were found to significantly enhance wafer level uniformity enabling fabrication of functional emitters within a few hundred [mu]m of the substrate boundary. The structures have been smoothed and rounded, making them suitable for use as ILIS, during a secondary etch process using electrolytes doped with nickel chloride to suppress transient effects. These doped solutions enabled a few [mu]m of material to be removed selectively from the porous surface while maintaining smooth features. These arrays have been mounted and aligned with electrostatic grids to demonstrate their emission capabilities. Propellant has been fed to the emitters by capillarity within the porous bulk and then extracted at potentials as low as 850 V. Beam currents exceeding several 100 [mu]A at both positive and negative polarities have been measured using both EMIIm and EMI-BF4 ionic liquid propellant. Two complete devices were tested yielding large beam currents and very high transmission fractions (- 88-100 %) from both attempts. We estimate that these devices can supply 10's of [mu]N of thrust at modest operating potentials, ~ 1.5 kV. with a specific impulse of roughly 2000-3000 s. When completely packaged, the thrusters measure 1.2 x 1.2 x 0.2 cm, weigh less than 1 g and require less than 0.65 W of operating power. These characteristics would be ideal for a small satellites where volume, mass and power are all at a premium, while the thrust levels would be sufficient to enable a variety of orbit variation and attitude control maneuvers. For example, applied to a CubeSat, this type of thruster system, including PPU, would occupy roughly 10 % of the spacecraft volume and mass while enabling de-orbiting from an 800 km altitude in roughly 100 days, compared with many years when left to decay naturally.

Investigation of Dual-mode Spacecraft Propulsion by Means of Ionic Liquids

Investigation of Dual-mode Spacecraft Propulsion by Means of Ionic Liquids
Author : Brian Russell Donius
Publisher : Unknown
Release Date : 2010
Category : Chemical equilibrium
Total pages :168
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"Analytical and numerical investigations of the performance of a series of potential dual-mode propulsion systems using ionic liquids are presented. Chemical bipropellant performance of select propellants is determined using NASAs Chemical Equilibrium with Applications. Comparison of the predicted specific impulse of ionic liquids with hydrazine and unsymmetrical dimethylhydrazine shows that the ionic liquid propellants have 3-12% lower specific impulse when paired with nitrogen tetroxide. However, when paired with hydroxylammonium nitrate, the specific impulse of the ionic liquids is 1-4% lower than that of hydrazine and unsymmetrical dimethylhydrazine paired with nitrogen tetroxide. Analytical investigation of an electrospray electric propulsion system shows that if ionic liquids are capable of operating in an almost purely ionic regime, they can provide very high specific impulse (~6000 sec). The predicted chemical and electric performance data are used in conjunction with system mass estimates to predict the system level performance of three dual-mode systems. Results indicate that the dual-mode systems are capable of producing higher change in velocity than traditional systems for any combination of chemical or electrical propulsion at the cost of time. Specifically, if 80% of the velocity change is accomplished by using electrical propulsion, a hydroxylammonium nitrate monopropellant electrospray system produced 190% more change in velocity than a traditional system consisting of a hydrazine monopropellant and xenon Hall effect thruster. This 190% increase in maximum velocity change comes at the cost 750% more time thrusting for the hydroxylammonium nitrate system"--Abstract, leaf iii.