Showing papers on "Spacecraft propulsion published in 2001"

Ion Propulsion Development Projects in US: Space Electric Rocket Test I to Deep Space 1

Abstract: The historical background and characteristics of the experimental flights of ion propulsion systems and the major ground-based technology demonstrations are reviewed. The results of the first successful ion engine flight in 1964, Space Electric Rocket Test (SERT) I, which demonstrated ion beam neutralization, are discussed along with the extended operation of SERT II starting in 1970. These results together with the technologies employed on the early cesium engine flights, the applications technology satellite series, and the ground-test demonstrations, have provided the evolutionary path for the development of xenon ion thruster component technologies, control systems, and power circuit implementations. In the 1997-1999 period, the communication satellite flights using ion engine systems and the Deep Space 1 flight confirmed that these auxiliary and primary propulsion systems have advanced to a high level of flight readiness.

A High Specific Impulse Two-Stage Hall Thruster with Plasma Lens Focusing

Abstract: * Presented as Paper IEPC-01-036 at the 27 International Electric Propulsion Conference, Pasadena, CA, 15-19 October, 2001. † Copyright © 2001 by the Electric Rocket Propulsion Society. All rights reserved. Recent interest in extending the performance envelope of Hall thrusters has spurred the development of several new thrusters built specifically for high specific impulse operation. The Plasmadynamics and Electric Propulsion Laboratory in conjunction with the NASA Glenn Research Center have built and successfully tested one of these thrusters in singleand two-stage operation at flow rates of 5-10 mg/s and total voltages of 300-600 V. The thruster employs a plasma lens field line topography that extends the plume focal length and an intermediate electrode composed of lanthanum hexaboride for operation in twostage mode. Performance and plume measurements have been conducted using a thrust stand and ion current density probes. The results of these measurements indicate that the plasma lens optimizes thruster performance and extends the plume focal length beyond one meter. Operation of the thruster in two-stage mode generally resulted in increased thrust above single-stage operation at the expense of efficiency.

Characterizing Vacuum Facility Backpressure Effects on the Performance of a Hall Thruster

Abstract: * Presented as Paper IEPC-01-045 at the 27 International Electric Propulsion Conference, Pasadena, CA, 15-19 October 2001. † Copyright © 2001 by the Electric Rocket Propulsion Society. All rights reserved. Vacuum facility backpressure is known to affect the performance of Hall thrusters through ingestion of background gases by the thruster. Ingested gases may be ionized and subsequently accelerated, artificially increasing the measured thrust. This study seeks to characterize how backpressure affects performance by changing the pumping speed of the facility, not by the commonly used technique of bleeding additional propellant in to the vacuum chamber. Performance is measured at xenon pumping speeds of 140,000 and 240,000 l/s, on a nominally 5 kW Hall thruster at voltages and flow rates of 300-600 V and 5-15 mg/s. At the highest pumping speed, performance data are collected while matching the discharge current observed at the lower pumping speed. Thrust, specific impulse, and efficiency are presented at both pumping rates. The anode mass flow at zero backpressure is calculated by extrapolation from the observed change in flow rate with pressure. On average, the correction is 4% of the anode flow rate at the maximum pumping speed, which is comparable to the uncertainty in the flow controllers.

Performance Measurements Using a Sub-Micronewton Resolution Thrust Stand ∗

Abstract: A sub-micronewton level thrust stand is used to evaluate the performance of five different microthrusters. Using a torsional balance design, the JPL microthrust stand (µTS) is capable of measuring steady-state thrust as low as 1 µN with sub-micronewton resolution and an impulse as low as 1 µNs with sub-micronewton-second resolution. The µTS can support a test object with a mass as large as 10 kg while supplying multiple diagnostic, power, and propellant lines to the thruster. The thrust or impulse is measured by monitoring the position of the thrust arm and using a calibrated dynamic model that will be presented in this paper. The performances of a vacuum arc thruster (VAT), gas-fed pulsed plasma thruster (GFPPT), indium field emission electric propulsion thruster (In-FEEP), vaporizing liquid microthruster (VLM), and micro-cold gas thruster have been measured using this test stand to validate its accuracy and precision.

Advanced Hall Electric Propulsion for Future In-space Transportation

Abstract: The Hall thruster is an electric propulsion device used for multiple in-space applications including orbit raising, on-orbit maneuvers, and de-orbit functions. These in-space propulsion functions are currently performed by toxic hydrazine monopropellant or hydrazine derivative/nitrogen tetroxide bi-propellant thrusters. The Hall thruster operates nominally in the 1500 sec specific impulse regime. It provides greater thrust to power than conventional gridded ion engines, thus reducing trip times and operational life when compared to that technology in Earth orbit applications. The technology in the far term, by adding a second acceleration stage, has shown promise of providing over 4000s Isp, the regime of the gridded ion engine and necessary for deep space applications. The Hall thruster system consists of three parts, the thruster, the power processor, and the propellant system. The technology is operational and commercially available at the 1.5 kW power level and 5 kW application is underway. NASA is looking toward 10 kW and eventually 50 kW-class engines for ambitious space transportation applications. The former allows launch vehicle step-down for GEO missions and demanding planetary missions such as Europa Lander, while the latter allows quick all-electric propulsion LEO to GEO transfers and non-nuclear transportation human Mars missions.

Thin Film Sensors for Surface Measurements

Abstract: Advanced thin film sensors that can provide accurate surface temperature, strain, and heat flux measurements have been developed at NASA Glenn Research Center. These sensors provide minimally intrusive characterization of advanced propulsion materials and components in hostile, high-temperature environments as well as validation of propulsion system design codes. The sensors are designed for applications on different material systems and engine components for testing in engine simulation facilities. Thin film thermocouples and strain gauges for the measurement of surface temperature and strain have been demonstrated on metals, ceramics and advanced ceramic-based composites of various component configurations. Test environments have included both air-breathing and space propulsion-based engine and burner rig environments at surface temperatures up to 1100 C and under high gas flow and pressure conditions. The technologies developed for these sensors as well as for a thin film heat flux gauge have been integrated into a single multifunctional gauge for the simultaneous real-time measurement of surface temperature, strain, and heat flux. This is the first step toward the development of smart sensors with integrated signal conditioning and high temperature electronics that would have the capability to provide feedback to the operating system in real-time. A description of the fabrication process for the thin film sensors and multifunctional gauge will be provided. In addition, the material systems on which the sensors have been demonstrated, the test facilities and the results of the tests to-date will be described. Finally, the results will be provided of the current effort to demonstrate the capabilities of the multifunctional gauge.

Future Spacecraft Propulsion Systems: Enabling Technologies for Space Exploration

Abstract: Overview- Our progress appears to be impeded- Commercial near-Earth space launcher: a perspective- Commercial near-Earth launcher: propulsion- Earth orbit on-orbit operations in near-Earth orbit, a necessary second step- Earth-Moon system: establishing a Solar System presence- Exploration of our Solar System- Stellar and interstellar precursor missions- View to the future and exploration of our Galaxy

Metal Matrix Composites for Liquid Rocket Engines

Abstract: This article presents an overview of current research and material requirements for metal-matrix composite (MMC) technologies being developed for liquid rocket engines (LRE) Developments in LRE technology for the US Air Force are being tracked and planned through the integrated high payoff rocket propulsion technologies program (IHPRPT) Current efforts and research requirements for three types of MMC systems are discussed: aluminum-, copper-, and nickel-matrix material systems Potential applications include turbopump housings, rotating machinery, and high-stiffness flanges and ductwork

Mini-magnetospheric plasma propulsion (M2P2): High speed propulsion sailing the solar wind

Abstract: Mini-Magnetospheric Plasma Propulsion (M2P2) seeks the creation of a magnetic wall or bubble (i.e. a magnetosphere) that will intercept the supersonic solar wind which is moving at 300–800 km/s. In so doing, a force of about 1 N will be exerted on the spacecraft by the spacecraft while only requiring a few mN of force to sustain the mini-magnetosphere. Equivalently, the incident solar wind power is about 1 MW while about 1 kW electrical power is required to sustain the system, with about 0.25–0.5 kg being expended per day. This nominal configuration utilizing only solar electric cells for power, the M2P2 will produce a magnetic barrier approximately 15–20 km in radius, which would accelerate a 70–140 kg payload to speeds of about 50–80 km/s. At this speed, missions to the heliopause and beyond can be achieved in under 10 yrs. Design characteristics for a prototype are also described.

Development and Testing of a 4500 Watt Flight Type Hall Thruster and Cathode

Abstract: * Presented as Paper IEPC-01-011 at the 27 International Electric Propulsion Conference, Pasadena, CA, 15-19 October, 2001. † Copyright  2001 by General Dynamics Space Propulsion Systems. Published by the Electric Rocket Propulsion Society with permission. A flight design BPT-4000 Hall thruster has been completed and two engineering model thrusters built and tested. The thruster development effort is part of a Lockheed Martin Space Systems Company (LMSSC) and General Dynamics Space Propulsion Systems (GD-SPS) funded program to develop a Hall Thruster Propulsion System (HTPS) for use on geosynchronous satellites. The thruster is a high performance, dual-mode capable design enabling both a high thrust/power and a high specific impulse.

The USAF Electric Propulsion Research Program

Abstract: : All overview of current electric propulsion research and development efforts within the United States Air Force is presented. The Air Force supports electric propulsion primarily through the Air Force Office of Scientific Research (AFOSR), the Air Force Research Laboratory (AFRL) and the AFOSR European Office of Aerospace Research and Development (BOARD). Overall direction for the programs comes from Air Force Space Command (AFSPC), with AFRL mission analysis used to define specific technological advances needed to meet AFSPC mission priorities. AFOSR funds basic research in electric propulsion throughout the country in both academia and industry. The AFRL Propulsion Directorate conducts electric propulsion efforts in basic research, engineering development, and space flight experiments. BOARD supports research at foreign laboratories that feeds directly into AFOSR and AFRL research programs. Current research efforts fall into 3 main categories defined loosely by the thruster power level. All three agencies are conducting research at the low-power regime (P 30 kW) is realizing increased emphasis.

Investigation of an Alternative Geometry Hybrid Rocket for Small Spacecraft Orbit Transfer

Abstract: : The following testing was carried out in support of the Surrey Space Centre (SSC) Alternative Geometry Hybrid Rocket research and development program. Although VFP testing was conducted in a low cost environment, the research program collected a wealth of valuable data with regard to this all-new hybrid rocket engine. The combustion efficiency is outstanding within the VFP. The scalability test has demonstrated that the VFP indeed scales well, providing high performance over the regimes tested as well as reliable, predictable, fuel liberation based upon the engine radius. The chamber pressure mapping did not reveal any pressure gradient across the diameter of the VFP rocket engine over the regimes tested. Flight propellant testing was promising in a number of areas. The VFP has demonstrated the ability to operate smoothly for long durations (up to 45 seconds tested), and return to within 1.5% of operational values upon relight (pulsed operations). It is likely that the engine may be burned near completion without the fear of solid fuel slivers blocking the rocket nozzle. The VFP test campaign provides solid evidence that the VFP is superior to conventional hybrid design in almost every respect and holds great promise for small spacecraft applications.

Thin film sensors for surface measurements [in aerospace simulation facilities]

Abstract: Advanced thin film sensors that can provide accurate surface temperature, strain, and heat flux measurements have been developed at NASA Glenn Research Center. These sensors provide minimally intrusive characterization of advanced propulsion materials and components in hostile, high-temperature environments as well as validation of propulsion system design codes. The sensors are designed for applications on different material systems and engine components for testing in engine simulation facilities. Thin film thermocouples and strain gauges for the measurement of surface temperature and strain have been demonstrated on metals, ceramics and advanced ceramic-based composites of various component configurations. Test environments have included both air-breathing and space propulsion-based engine and burner rig environments at surface temperatures up to 1100/spl deg/C and under high gas flow and pressure conditions. The technologies developed for these sensors as well as for a thin film heat flux gauge have been integrated into a single multifunctional gauge for the simultaneous real-time measurement of surface temperature, strain, and heat flux. This is the first step toward the development of smart sensors with integrated signal conditioning and high temperature electronics that would have the capability to provide feedback to the operating system in real-time. A description of the fabrication process for the thin film sensors and multifunctional gauge is provided. In addition, the material systems on which the sensors have been demonstrated, the test facilities and the results of the tests to-date are described. Finally, the results are provided of the current effort to demonstrate the capabilities of the multifunctional gauge.

The Development and Qualification of a 4.5 kW Hall Thruster Propulsion System for GEO Satellite Applications

Abstract: * Presented as Paper IEPC-01-010 at the 27 International Electric Propulsion Conference, Pasadena, CA, 15-19 October, 2001. † Copyright  2001 by General Dynamics SPS. Published by the Electric Rocket Propulsion Society with permission. This paper summarizes the current program at General Dynamics Space Propulsion Systems (formerly Primex Aerospace Company) and Lockheed Martin to develop and qualify a 4.5 kW Hall Thruster Propulsion Subsystem. This subsystem is being developed to support geo-synchronous satellite applications. This paper describes the mission application, performance, program plan and current status of the development program. The overall system, including the Power Processor Unit, the Xenon Flowrate Controller, the Hall Thruster and Cathode and the Xenon Feed System is also described.

High Voltage Solar Arrays for a Direct Drive Hall Effect Propulsion System

Abstract: * Presented as Paper IEPC-01-000 at the 27 International Electric Propulsion Conference, Pasadena, CA, 15-19 October, 2001. † Copyright © 2001 by the Electric Rocket Propulsion Society. All rights reserved. Now at JPL Gary A. Jongeward, Ira Katz, Ioannis G. Mikellides Science Applications International Corporation 10260 Campus Point Dr, MS X-1 San Diego, CA 92121 (858) 826-1624 Gary.A.Jongeward@saic.com Melvin. R. Carruth NASA/MSFC (256) 544-7647 ralph.carruth@msfc.nasa.gov

Indium FEEP Thruster Beam Diagnostics, Analysis and Simulation

Abstract: Space Propulsion Austrian Research Centers Seibersdorf, A-2444 Seibersdorf, Austria Indium FEEP thrusters based on space proven needle type liquid metal ion emitters are presently considered for a variety of missions which need ultraprecise drag-free capabilities such as LISA, SMART-2, GOCE or DIVA. One of the requirements of these missions is to investigate possible thrust vector variations, therefore the ion beam profile needs to be measured. A dedicated beam diagnostic assembly consisting of a wire and a Langmuir probe was installed in the laboratory to investigate single and multi-emitter configurations in the full thrust level of 1 – 100 µN per emitter. The derived thrust coefficient, thrust vector angle and beam divergence were within the expected limits. Multi-emitter configurations did not show any significant ion beam interaction. This analysis was supported by numerical simulations that could obtain similar results for emission in a pure vacuum environment without ambient electrons that can provide neutralization.

PIC Simulation of High Specific Impulse Hall Effect Thruster

Abstract: * Presented as Paper IEPC-01-037 at the 27 International Electric Propulsion Conference, Pasadena, CA, 15-19 October, 2001. † Copyright © 2001 by the Electric Rocket Propulsion Society. All rights reserved. A previously developed computational method, which treats all particles in the plasma kinetically, is applied for a study of the effects of increasing the voltage beyond the wellestablished range around 300V (for Xenon). Although some additional code development is still necessary to improve absolute accuracy, several important trends and effects are identified through these calculations. If the magnetic field is kept constant, the overall anode efficiency is found to increase at first, peak at about 600V, then decrease at higher voltages. On the other hand, if B is optimized at each voltage, the efficiency increases continuously with V. The detailed physics behind this behavior are identified. The secondion fraction (Xe/Xe) increases rapidly at first, but nearly saturates with further voltage increases. The electron temperature approaches proportionality with voltage, since metallic walls (TAL style) were assumed, and secondary electron emission was ignored.

Integrated System Test of an Airbreathing Rocket (ISTAR)

Abstract: Rocket Based Combined Cycle (RBCC) propulsion system development and ground test is being conducted as part of the NASA Marshall Space Flight Center Integrated System Test of an Airbreathing Rocket (ISTAR) program. Rocketdyne, Aerojet and Pratt & Whitney have teamed as the Rocket Based Combined Cycle Consortium (RBC3) to work the propulsion system development. Each company offered unique RBCC propulsion concepts as candidates for the ISTAR propulsion system. A team of engine contractor, vehicle contractor and NASA representatives reviewed the concepts proposed by each company, reviewed the available data and selected the Aerojet RBCC propulsion system concept as the team propulsion system baseline for the ISTAR program. The ISTAR program is currently in a "Jumpstart" phase for development of the engine system leading to ground test of a thermally and power balanced RBCC propulsion system at Stennis Space Center in 2005. A parallel flight test demonstration of this propulsion system is anticipated to lead to first flight in the 2007 timeframe.

Phase change material thermal control for electric propulsion

Abstract: Systems and methods that employ a phase change material to provide thermal control of electric propulsion devices (thrusters). A spacecraft is configured to have an electric propulsion thruster. The electric propulsion thruster is surrounded with a phase change material. Suitable phase change materials include high-density polyethylene (HDPE), waxes, paraffin materials, and eutectic salts. The spacecraft is launched into orbit. The electric propulsion thruster is fired for a predetermined period of time. Heat generated by the electric propulsion thruster is absorbed and stored in the phase change material while the thruster is firing. The stored heat is dissipated into space after the thruster has stopped firing.

Experimental Investigation of a Hall Thruster Internal Magnetic Field Topography

Abstract: * Presented as Paper IEPC-01-030 at the 27 International Electric Propulsion Conference, Pasadena, CA, 15-19 October 2001. † Copyright © 2001 by Peter Y Peterson. Published by the Electric Rocket Propulsion Society with permission. ‡ Graduate Student, Aerospace Engineering. § Associate Professor, Aerospace Engineering and Applied Physics. ** Former Graduate Student, Aerospace Engineering, Currently a Research Scientist at Edwards Air Force Base

Engineering the Future of Full-Scale Propulsion Testing

Abstract: Year 2000 has been an active one for rocket propulsion testing at the NASA John C. Stennis Space Center. This paper highlights several major test facilites for large-scale propulsion devices, and summarizes the varied nature of the recent test projects conducted at the Stennis Space Center (SSC) such as the X-33 Aerospike Engine, Ultra Low Cost Engine (ULCE) thrust chamber program, and the Hybrid Sounding Rocket (HYSR) program. Further, an overview of relevant engineering capabilities and technology challenges in conducting full-scale propulsion testing are outlined.

Plume Characteristics of the Indium Needle Emitter (InFEEP) Thruster

Abstract: * Presented as Paper IEPC-01-291 at the 27 International Electric Propulsion Conference, Pasadena, CA, 15-19 October, 2001. † Copyright © 2001 by ESA. Published by the Electric Rocket Propulsion Society with permission Characterization test campaigns involving ion beam analysis and overall performance evaluation are being performed in the dedicated FEEP test facility at the ESA Electric Propulsion Laboratory on the European Field Emission Thrusters developed under ESA funding. These are the ARCS Indium Liquid Metal Ion Source (LMIS, also referred as InFEEP) and the Alta FEEP. This paper presents part of the achieved results on the first series of tests carried out by ESA on two ARCS InFEEP modules.

Advanced Pulsed Plasma Thruster Electrical Components

Abstract: An overview of the National Aeronautics and Space Administration and industry project to advance the state of pulsed plasma thruster (PPT) electrical component technology is presented. Following the successful development of the PPT for the Earth Observing-1 mission, the NASA John H. Glenn Research Center contracted with industry to develop advanced energy storage, discharge initiation and power processing components to meet future mission PPT needs. Unison Industries, with support from General Dynamics Space Propulsion Systems and CU Aerospace, developed breadboard level components that provide projected benefits in low mass, long life, low impulse bit variability and operational flexibility. This paper describes the government/industry process of PPT architecture definition, technology selection, and component design. Key features of the resulting designs are related to future mission capabilities.

SailBeam: Space propulsion by macroscopic sail-type projectiles

Abstract: Traditional beam-driven (laser, microwave, or particle beam) sail concepts for interstellar or interstellar precursor missions envision a single large sail pushed to its final velocity by a single beam source. This paper considers the possibility of using many macroscopic (microgram to gram) laser-sail-driven projectiles (microsails) to propel a larger mission vehicle via momentum transfer. Scaling relationships indicate that absolute performance, as measured by beam source parameters required to accelerate a fixed vehicle mass to a given velocity, can be improved by several orders of magnitude relative to a traditional sail. Also, beam pointing requirements are relaxed, and guidance and stability problems may be reduced. The concept is also applicable to high-velocity solar-sail missions, with possible substantial benefits to mission design.

Sub-Kilowatt Radioisotope Electric Propulsion for Outer Solar System Exploration

Abstract: An activity has begun to perform a multi-mission trajectory/systems analysis study that examines the combined benefit of three key technology development areas The three key technologies leveraged in the study include sub-kilowatt ion propulsion, Stirling radioisotope power systems and microelectronics/lightweight spacecraft bus technologies This study is being performed jointly by NASA Glenn Research Center and the Applied Physics Laboratory to leverage their combined areas of expertise in advanced power/propulsion and spacecraft design Missions examined in this study include missions to outer planets' moons, such as Europa, Titan, and Triton, and a Comet Nucleus Sample Return mission Additional information is contained in the original extended abstract

A Xenon Flowrate Controller for Hall Current Thruster Applications 1

Abstract: As Electric Propulsion (EP) becomes more prevalent within the spacecraft propulsion community, there is an increased need for greater control of the xenon flow. Moog has concluded a Research and Development project that resulted in the Moog Proportional Flow Control Valve (PFCV), which can be used to throttle the flow of xenon over a wide range of inlet pressures and flow rates. Refer to

An Overview of Computational Fluid Dynamics for Application to Advanced Propulsion Systems

Abstract: Validation of computational fluid dynamics (CFD) codes appropriate for subsonic through hypersonic applications requires careful consideration of the physical processes encountered in these flight regimes and detailed comparisons with quality experimental datasets that simulate these processes. The primary objective of this study is to overview the state of the practice of current CFD tools used for advanced propulsion system design and analysis. In addition, this paper provides an overview of the validation requirements of CFD codes applicable to advanced propulsion systems, and identifies the enabling technology requirements in the development of the next generation of CFD tools. Several major propulsion systems such as rockets, engines, plumes, scramjets, and gas turbine and pulse detonation engines are discussed. The physical processes that dominate combustion phenomenology are described for various engine types. An assessment of the state of the art of numerical approximations used to simulate the flow-field effects of these phenomena is provided. Computational issues related to numerical algorithms and the implementation of physical models are also addressed. During the last several years, most U.S. industries have witnessed major changes in their business environment. These changes include significantly curtailed government spending, reduced corporate product development budgets, and the transition from a national to a global economy. In view of these changes, U.S. industries have been forced to critically review their basic engineering and manufacturing processes and identify potential cost-saving initiatives to maintain their business share and broaden their market base. The traditional design and analysis practices rely heavily on costly full-scale prototype development and testing. Introduction of high-fidelity design and analysis tools such as computational fluid dynamics (CFD) early in the product development cycle was identified as one way to alleviate the testing costs and develop products "better, faster, and cheaper." In the design of advanced propulsion systems, computational modeling plays a major role in defining the required performance over the flight envelope, as well as in testing the sensitivity of the design to the various modes of operation (e.g., afterburner, rocket, ramjet, and scramjet). Computational modeling techniques, complemented with select ground and flight testing, are expected to be the engineering approach of choice in the development of new Air Force and NASA space propulsion programs. Therefore, increased emphasis is placed on developing and applying CFD models to simulate the flow-field environments and performance of advanced propulsion systems. This places a premium on the development of the next generation of computational tools so that this can be used effectively and reliably in a design environment by non-CFD specialists. Experience gained from the use of current research-oriented CFD models is essential to guide the successful development of engineering application tools. Since the new approaches will rely less on testing and more on CFD results, a careful strategy is needed to ensure that the models are appropriate for the phenomena being simulated and that they are applied in a timely and efficient manner. An approach to assessing the CFD model involves the methodology described as follows. Currently, available CFD codes are validated by careful comparison with measurements. The validation determines the range of validity of the model and identifies improvements needed in the physical approximations and numerical algorithms.

Specifics of Small Satellite Propulsion: Part 1

Abstract: Small satellite propulsion is a subject of unique constraints and requirements. Based on University of Surrey experience in small satellite building and operation, these features are listed and explained. Available volume is often identified as the most severe constraint for a small satellite with power and cost being the other two major constraints. Mass is often only of secondary importance for small satellites. Propulsion dry mass fraction for a spacecraft grows upon the system scaling-down. For small spacecraft propulsion fraction can easily exceed 85%. In such a case, a combination of independent systems for multifunctional propulsion mission scenarios would aggravate the situation. Moreover, specific impulse is not a factor reflecting small satellite propulsion system performance since spacecraft velocity change is also a function of propulsion dry mass fraction. New conceptual and design solutions are suggested for small satellite propulsion with respect to its specific constraints and requirements. Features of future advanced, low-cost propulsion system for small satellite are described.