A short review of the status of electric propulsion (EP) is presented to serve as an introduction to the more specialized technical papers also appearing in this Special Issue (Journal of Propulsion and Power, Vol. 14, No. 5, Sept. –Oct. 1998). The principles of operation and the several types of thrusters that are either operational or in advanced development are discussed  rst, followed by some considerations on the necessary power sources. A few prototypical missions are then described to highlight the operational peculiarities of EP, including spacecraft interactions. We conclude with a historical summary of the accumulated  ight experience using this technology.

/pdf/spacecraft-electric-propulsion-an-overview-cprtlkf4bz.pdf

Spacecraft Electric Propulsion—An Overview

The 35-year history of the development and application of the pulsed plasma thruster (PPT) is reviewed. The PPT operates by creating a pulsed, high-current discharge across the exposed surface of a solid insulator, such as a Tee on t bar. The arc discharge ablates material from this surface, thereby providing propellant that is ionized, heated, and accelerated to high speed. Typically, the current pulse lasts for a few microseconds, driven by a capacitor that is charged and discharged approximately once per second. Exhaust speeds range from 3 to 50 km/s, depending on the details of the PPT design. We review the basic physics and types of PPTs, and discuss the performance of e ight and laboratory versions, with special attention to velocity and plume measurements. We also present the status of PPT theory and modeling, with emphasis on mass evolution and plasma acceleration, and describe recent variations on PPT operation, laboratory thrust measurement techniques, and the separate components of PPT efe ciency.

Pulsed Plasma Thruster

This contribution presents a short review of electric propulsion (EP) technologies for satellites and spacecraft. Electric thrusters, also termed ion or plasma thrusters, deliver a low thrust level compared to their chemical counterparts, but they offer significant advantages for in-space propulsion as energy is uncoupled to the propellant, therefore allowing for large energy densities. Although the development of EP goes back to the 1960s, the technology potential has just begun to be fully exploited because of the increase in the available power aboard spacecraft, as demonstrated by the very recent appearance of all-electric communication satellites. This article first describes the fundamentals of EP: momentum conservation and the ideal rocket equation, specific impulse and thrust, figures of merit and a comparison with chemical propulsion. Subsequently, the influence of the power source type and characteristics on the mission profile is discussed. Plasma thrusters are classically grouped into three categories according to the thrust generation process: electrothermal, electrostatic and electromagnetic devices. The three groups, along with the associated plasma discharge and energy transfer mechanisms, are presented via a discussion of long-standing technologies like arcjet thrusters, magnetoplasmadynamic thrusters, pulsed plasma thrusters and ion engines, as well as Hall thrusters and variants. More advanced concepts and new approaches for performance improvement are discussed afterwards: magnetic shielding and wall-less configurations, negative ion thrusters and plasma acceleration with a magnetic nozzle. Finally, various alternative propellant options are analyzed and possible research paths for the near future are examined.

http://iopscience.iop.org/article/10.1088/0963-0252/25/3/033002/pdf

Electric propulsion for satellites and spacecraft: established technologies and novel approaches

Sputtering yields for different ions and materials at low ion energies have a similar energy dependence. Due to this similarity, yield data can be characterized by a normalized energy function and two parameters for each ion target combination. One of these parameters is the threshold energy. An energy scaling can be based on this parameter. The other parameter is a multiplication factor. Both parameters depend mainly on the ion and target mass M1 and M2 and on the surface binding energy EB. An analytic expression for the normalized functions and both the parameters is given. This empirical relation also allows an estimate of unknown sputtering data, if M1, M2, and EB are noted. A physical interpretation of the empirical relation is given for the case M1≪M2, as in this case special collision processes which dominate the sputtering can be identified.

An Analytical Formula and Important Parameters for Low-Energy Ion Sputtering: Erratum of the Publication in Journal of Applied Physics Vol. 51, S. 2861 ff, (1980)

The technological and commercial expansion of electric propulsion

Solar electric propulsion technology is currently being used for geostationary satellite station keeping. Analyses show that electric propulsion technologies can be used to obtain additional increases in payload mass by using them to perform part of the orbit transfer. Three electric propulsion technologies are examined at two power levels for geostationary insertion of an Atlas IIAS class spacecraft. The onboard chemical propulsion apogee engine fuel is reduced in this analysis to allow the use of electric propulsion. A numerical optimizer is used to determine the chemical burns that will minimize the electric propulsion transfer times. For a 1550-kg Atlas IIAS class payload, increases in net mass (geostationary satellite mass less wet propulsion system mass) of 150-800 kg are enabled by using electric propulsion for station keeping, advanced chemical engines for part of the transfer, and electric propulsion for the remainder of the transfer. Trip times are between one and four months.

Advanced Propulsion for Geostationary Orbit Insertion and North-South Station Keeping

Advanced chemical and low power electric propulsion offer attractive options for small satellite propulsion. Applications include orbit raising, orbit maintenance, attitude control, repositioning, and deorbit of both Earth-space and planetary spacecraft. Potential propulsion technologies for these functions include high pressure Ir/Re bipropellant engines, very low power arcjets, Hall thrusters, and pulsed plasma thrusters, all of which have been shown to operate in manners consistent with currently planned small satellites. Mission analyses show that insertion of advanced propulsion technologies enables and/or greatly enhances many planned small satellite missions. Examples of commercial, DoD, and NASA missions are provided to illustrate the potential benefits of using advanced propulsion options on small satellites.

Small satellite propulsion options

Stationary Plasma Thrusters (SPT's) are being investigated for application to a variety of near-term missions This paper presents the results of a preliminary study of the thruster plume characteristics which are needed to assess spacecraft integration requirements Langmuir probes, planar probes, Faraday cups, and a retarding potential analyzer were used to measure plume properties For the design operating voltage of 300 V the centerline electron density was found to decrease from approximately 18 x 10 exp 17 cubic meters at a distance of 03 m to 18 X 10 exp 14 cubic meters at a distance of 4 m from the thruster The electron temperature over the same region was between 17 and 35 eV Ion current density measurements showed that the plume was sharply peaked, dropping by a factor of 26 within 22 degrees of centerline The ion energy 4 m from the thruster and 15 degrees off-centerline was approximately 270 V The thruster cathode flow rate and facility pressure were found to strongly affect the plume properties In addition to the plume measurements, the data from the various probe types were used to assess the impact of probe design criteria

/pdf/stationary-plasma-thruster-plume-characteristics-3isyy7z15h.pdf

Stationary Plasma Thruster Plume Characteristics

Integration of pulsed plasma thruster (PPT) onboard spacecraft requires the evaluation of potential plume/spacecraft interactions. Important e ndings are summarized of our experimental and modeling plume investigations of rectangular-geometry Tee on ® PPTs. Initial studies of the Lincoln Experimental Satellite 8 /9 PPT plume used time-of-e ight analysis of single langmuir probe data and found two ion populations with approximately 30 and 60 km /s, respectively. A residual gas analyzer identie ed C, F, C xFy, and various thruster materials. Fast ionization gauges detected the presence of slow neutral particles up to 1 ms after the end of the discharge. Subsequent studies used triple langmuir probes and obtained electron temperature and density in the plume of a laboratory PPT operating at 5, 20, and 40 J. Plume properties showed large angular density variation on the perpendicular to the electrodes plane but small variation on the parallel plane. Electron density and temperature were found to decrease with increasing radial distance from the Tee on surface. Time-average temperatures were between 1 and 3 eV. Time-averageelectron density increased with increasing dischargeenergy and are in therange between 10 19 and 2 ££ 10 20 m i 3 for 5 J, 6 ££ 10 20 to 10 21 m i 3 for 20 J, and 2 ££ 10 20 to 1.4 ££ 10 21 m i 3 for 40 J. The PPT plume model is based on a hybrid (particle‐e uid) methodology. Neutrals and ions were modeled with a combination of the direct simulation Monte Carlo and a hybrid-particle-in-cell method. Electrons were modeled as a massless e uid with a momentum equation that includes collisional contributions from ions and neutrals and an energy equation. Simulations of the laboratory PPT operating at discharge energies of 5, 20, and 40 J showed the expansion of the neutral and ion components of the plume during a pulse, the generation of low-energy ions and high-energy neutrals dueto chargeexchange reactions, and the generation ofbacke ow. Numerical predictions showed good quantitative agreements with data.

Roger M. Myers

Papers

The technological and commercial expansion of electric propulsion

Advanced Propulsion for Geostationary Orbit Insertion and North-South Station Keeping

Small satellite propulsion options

Stationary Plasma Thruster Plume Characteristics

Experimental Investigations and Numerical Modeling of Pulsed Plasma Thruster Plumes