Thrusters that exploit vacuum arc discharges to produce high-velocity plasma jets directly or as sources of plasma that is subsequently accelerated electrostatically have been proposed or are currently under development. Vacuum arc discharges exhibit certain regularities in their behavior that allow the performance of these thrusters to be described by simple semiempirical models. Empirical data on the current density distribution, charge state and velocity of ions created in vacuum arc discharges, and the total cathode mass loss rate are used to develop expressions for the expected thrust and specific impulse as a function of thruster geometry. Thruster electrical efficiency and thrust-to-power ratio are calculated based on measurements of the burning voltage for given thruster operating parameters. Estimates of achievable thruster performance for a wide range of cathode materials are presented. This analysis suggests that thrusters using vacuum arc sources can be operated efficiently with a range of propellant options that gives great flexibility in specific impulse. In addition, the efficiency of plasma production in these devices appears to be largely independent of scale because the metal vapor is ionized within tens of micrometers of the cathode electron emission sites, so this approach is well suited for micropropulsion.

The accurate, direct measurement of thrust or impulse is one of the most critical elements of electric thruster characterization, and it is one of the most difficult measurements to make. This paper summarizes recommended practices for the design, calibration, and operation of pendulum thrust stands, which are widely recognized as the best approach for measuring micronewton-to millinewton-level thrust and micronewton-per-second-level impulse bits. The fundamentals of pendulum thrust stand operation are reviewed, along with the implementation of hanging pendulum, inverted pendulum, and torsional balance configurations. The methods of calibration and recommendations for calibration processes are presented. Sources of error are identified, and methods for data processing and uncertainty analysis are discussed. This review is intended to be the first step toward a recommended practices document to help the community produce high-quality thrust measurements.

We present a new gas-fed pulsed plasma thruster performance database and use it to investigate the electromagnetic nature of the acceleration process by identifying trends in the dependencies of the performance parameters on the mass bit, energy, and capacitance. It was previously observed[1] that performance measurements for PPTs operating at low mass bits (< 2 µg per pulse) can be influenced by contamination from diffusion pump oil. Consequently, facility renovations and a new impulse measurement protocol using liquid nitrogen cooled baffles to reduce diffusion pump oil backstreaming were implemented and are discussed. A portion of the database has been duplicated in a cryo-pumped facility at NASA-JPL to verify that the contamination effects have been removed. The intermittent simultaneous operation of the discharge initiation plugs, that was previously observed to degrade the performance, was solved with a new electric circuit and the resulting discharge symmetry verified using a fast-framing camera. The nature of the acceleration process is examined by comparing trends in the performance database with the theoret-

We present a characterization of the performance of a recently developed gas-fed pulsed plasma thruster (GF-PPT) at low discharge energies (≤5 J). The impulsive thrust measurements were made using EPPDyL's high-accuracy interferometric microthrust stand. The thruster is best suited for small satellite applications and is operated in an unsteady pulsed mode (3 µs/pulse). It is the result of a series of design iterations aimed at achieving the highest thrust efficiencies for unsteady electromagnetic acceleration at low discharge energies. The use of advanced nonlinear magnetic switching technology, which insured a total system inductance of 3-4 nH, combined with an electrode geometry and radial gas injection that favor low profile losses, yielded a total efficiency of 50% at 5 J with argon (at an impulse bit of 32 µNs and a mass bit of .2 µg/shot). This is the highest measured efficiency ever reported for a PPT at this low energy level. Moreover, the low mass utilization efficiency problem that plagued previous gas-fed pulsed

The scaling laws of pulsed plasma thrusters operating in the predominantly electromagnetic acceleration mode (EM-PPT) are investigated theoretically and experimentally using gas-fed pulsed plasma thrusters. A fundamental characteristic velocity that depends on the inductance per unit length and the square root of the capacitance to the initial inductance ratio is identified. An analytical model of the discharge current predicts scaling laws in which the propulsive efficiency is proportional to the EM-PPT performance scaling number, defined here as the ratio of the exhaust velocity to the EM-PPT characteristic velocity. The importance of the effective plasma resistance in improving the propulsive performance is shown. To test the validity of the predicted scaling relations, the performance of two gas-fed pulsed plasma thruster designs (one with coaxial electrodes and the other with parallel-plate electrodes), was measured under 70 different operating conditions using an argon plasma. The measurements demonstrate that the impulse bit scales linearly with the integral of the square of the discharge current as expected for an electromagnetic accelerator. The measured performance scaling is shown to be in good agreement with the theoretically predicted scaling. Normalizing the exhaust velocity and the impulse-to-energy ratio by the EM-PPT characteristic velocity collapses almost all the measured data onto single curves that uphold the general validity of these scaling laws.

An optical interferometric proximeter system (IPS) for measuring thrust and impulse bit of pulsed electric thrusters was developed. Unlike existing thrust stands, the IPS-based thrust stand offers the advantage of a single system that can yield electromagnetic interference-free, high accuracy (&lt;2% error) thrust measurements within a very wide range of impulses (100 μN s to above 10 N s) covering the impulse range of all known pulsed plasma thrusters. In addition to pulsed thrusters, the IPS is theoretically shown to be capable of measuring steady-state thrust values as low as 20 μN for microthrusters such as the field emission electric propulsion thruster. The IPS-based thrust stand relies on measuring the dynamic response of a swinging arm using a two-sensor laser interferometer with 10 nm position accuracy. The wide application of the thrust stand is demonstrated with thrust measurements of an ablative pulsed plasma thruster and a quasi-steady magnetoplasmadynamic thruster.

Electron energy distribution function in a Hall-type thruster (abstract) Rev. Sci. Instrum. 71, 731 (2000)A new type of vacuum arc thruster in combination with an innovative power processing unit ͑PPU͒ has been developed that promises to be a high efficiency ͑ϳ15%͒, low mass ͑ϳ100 g͒ propulsion system for micro-and nanosatellites. This thruster accelerates a plasma that consists almost exclusively of ions of the cathode material and has been operated with a wide variety of cathodes. The streaming velocity of the plasma exhaust varies with cathode material, from a low of 11 km/s for Ti up to 30 km/s for Al, with a corresponding range of specific impulse from 1100 s for Ta to 3000 s for Al. Initiation of the arc requires only a few hundred volts due to an innovative ''triggerless'' approach in which a conductive layer between the cathode and the anode produces the initial charge carriers needed for plasma production. The initial starting voltage spike as well as the energy to operate the vacuum arc are generated by a low mass ͑Ͻ300 g͒ inductive energy storage PPU which is controlled using ϩ5 V level signals. The thrust-to-power ratio has been estimated to reach up to Ϸ20 N/W. The vacuum arc thruster was tested at the Jet Propulsion Laboratory using W as cathode material. Experimental results are within 65% of the estimated values.

The Space Technology 7 Disturbance Reduction System (ST7-DRS) is a NASA technology demonstration payload that operated from January 2016 through July of 2017 on the European Space Agency's LISA Pathfinder spacecraft. The joint goal of the NASA and ESA missions was to validate key technologies for a future space-based gravitational wave observatory targeting the source-rich milliHertz band. The two primary components of ST7-DRS are a micropropulsion system based on colloidal micro-Newton thrusters (CMNTs) and a control system that simultaneously controls the attitude and position of the spacecraft and the two free-flying test masses (TMs). This paper presents our main experimental results and summarizes the overall the performance of the CMNTs and control laws. We find that the CMNT performance to be consistent with pre-flight predictions, with a measured system thrust noise on the order of 100 nN/ √ Hz in the 1 mHz ≤ f ≤ 30 mHz band. The control system maintained the TM-spacecraft separation with an RMS error of less than 2 nm and a noise spectral density of less than 3 nm/ √ Hz in the same band. Thruster calibration measurements yield thrust values consistent with the performance model and ground-based thrust-stand measurements, to within a few percent. We also report a differential acceleration noise between the two test masses with a spectral density of roughly 3 fm/s 2 / √ Hz in the 1 mHz ≤ f ≤ 30 mHz band, slightly less than twice as large as the best performance reported with the baseline LISA Pathfinder configuration and below the current requirements for the Laser Interferometer Space Antenna (LISA) mission.