High Power Electric Propulsion Systems Research Articles

Nature-Inspired Concepts for High-Power Electric Propulsion Systems

The development of new space missions and the growing interest in space exploration have created an urgent need to develop high-thrust propulsion systems capable of propelling spacecraft far beyond the Earth and the solar system for long periods. Electric propulsion can potentially enable space missions to reach speeds thousands of times greater than conventional high-thrust chemical rockets. However, high speed comes at the cost of low power-to-thrust efficiency when considering propulsion systems as a whole, including the power generation system, transmission lines, and thrusters, which prevents high thrust from being achieved with any conceivable power system, resulting in long acceleration times. In addition, modern electric propulsion systems rely on external power sources that suffer significant power transfer losses at the high power levels required for high thrust levels. In addition, modern electric propulsion systems suffer from a number of critical physical and engineering problems that affect thrust levels and longevity. In addition, modern electric propulsion systems do not follow the principles of generation and acceleration of plasma flow that can be observed in space and potentially borrowed for artificial applications. This paper discusses several promising electrodeless plasma thruster concepts for high-power, high-thrust electric propulsion systems based on a combined power source/power converter/thruster architecture. These concepts have the potential to overcome modern limitations of high-power electric propulsion systems and enable new outer space missions that would not be possible with conventional thrusters.

Development of inverted pendulum thrust stand with spring-shaped wire for high power electric thrusters.

Pendulum thrust stands are used to measure the thrust of electric propulsion systems for spacecraft. A thruster is mounted on a pendulum and operated, and the pendulum displacement due to thrust is measured. In this type of measurement, the pendulum is also affected by nonlinear tensions due to wiring and piping that deteriorate the accuracy of the measurement. This influence cannot be ignored in high power electric propulsion systems because complicated piping and thick wirings are required. Therefore, to reduce the influence of tension due to wires and tubes, we developed an inverted pendulum-type thrust stand with pipes and wirings as springs. In this paper, we first derive the design guidelines for spring-shaped wires; the necessary conditions for sensitivity, responsivity, spring shape, and electric wire were formulated. Next, a thrust stand was designed and fabricated based on these guidelines, and the performance of the stand was evaluated through calibration and thrust measurements using a 1 kW-class magneto-plasma-dynamics thruster. The sensitivity of the thrust stand was 17 mN/V, the normalized standard deviation of the variation of the measured values owing to the structure of the thrust stand was 1.8 × 10-3, and the thermal drift during the long-time operation was ∼4.5 × 10-3 mN/s.

Development of a 50-A heaterless hollow cathode for electric thrusters.

Hollow cathodes in electric thrusters normally use an external heater to raise the thermionic electron emitter to emission temperatures. These heaters are a potential single-point failure in the thruster and add a separate power supply to the power processing unit. Heaterless hollow cathodes are attractive for their compact size and potential higher reliability but have only been reliably demonstrated to date in small hollow cathodes capable of discharge currents below around 5A. A new heaterless LaB6 hollow cathode has been developed that is capable of discharge currents from 5 to 50A. The cathode configuration extends the gas feed tube at cathode potential part way into the emitting insert region of the cathode. A high-voltage Paschen discharge is struck from the tube to the keeper that heats the tube tip, which then efficiently heats the insert by radiation. This configuration eliminates the arcing observed in prior large heaterless designs that coupled the high-voltage Paschen discharge to the orifice plate or the insert itself. Discharge current-voltage characteristics show that the presence of the tube does not significantly perturb the insert-region plasma. Startup uses a simple 3minignition procedure, and voltage traces of the keeper discharge reveal that much of the present tube-radiator's 100-to-150W heating power comes from an intermediate thermionic discharge sustained by the tube during the transition between the Paschen discharge and LaB6 insert thermionic regime. This novel heating mechanism enables an unprecedented class of higher-current heaterless hollow cathodes for the next generation of high-power electric propulsion systems.

Структуры систем электродвижения большой мощности

Object and purpose of research. This paper discusses structures of high-power electric propulsion systems for ships. The purpose was to give a summary of design solutions made in development of these systems. Materials and methods. This paper relies on academic and technical data, as well on the long-term author’s experience in marine electric propulsion R&Ds. The solution suggested by the author is based on the comparative analysis of design solutions adopted in the development of structures for high-power marine electric power and propulsion systems. Main results. Summary on design solutions for high-power electric propulsion systems of such ships as icebreakers, oil tankers, LNGCs and cruise liners. Conclusion. Results obtained by author were used in the design of the electric propulsion system of the «Lider» nuclear icebreaker and further could be used in design of Arctic vessels.

Influences of cathode position on operations and thrust performance of side by side Hall thruster system

High-power electric propulsion systems are promising for future large-scale space missions. To develop a high-power Hall thruster cluster system, researchers at the Gifu University have been investigating the effects of plume interference using 500 W class side-by-side (SBS) systems having two magnetic-layer heads. In this study, to investigate the dependencies of discharge characteristics and thrust performance on the axial position of a cathode, we measured the discharge currents and the thrusts for various axial positions. The optimum cathode position for the thrust performance was examined using the results. As the cathode shifted downstream, thrust decreased more than 10% for each operation. The results also revealed that at a high voltage in the SBS operations, the thrust performance increases for shifting the cathode close to the head end.

Low-thrust roundtrip trajectories to Mars with one-synodic-period repeat time

Cycler trajectories—both ballistic and powered—are reported in the literature in which there are two-vehicle, three-vehicle, and four-vehicle cases. Such trajectories permit the installation of cycler vehicles which provide safe and comfortable living conditions for human space travel between Earth and Mars during every synodic opportunity. The question the present paper answers is a logical, obvious one: Does a single-vehicle, one-synodic-period cycler exist? The answer is yes: such a trajectory can be flown—but only with a high-power electric propulsion system. In our example, it is found that “stopover” trajectories that spend 30 days in orbit about Earth and 30 days about Mars, and return astronauts to Earth in one synodic period require a 90-t power generator with a power level of 11 MWe. Fortuitously, and in lieu of using chemical propulsion, the high power level of the electric propulsion system would also be effective in hauling the cargo payload via a spiral trajectory about the Earth. But because one synodic period is not enough for the cycler vehicle to fly both the interplanetary trajectories and the Earth-spiral trajectories, we suggest developing two nuclear power generators, which could alternate flying the interplanetary trajectories and the Earth-spiral trajectories. Once these power generators are launched and begin operating in space, the mass requirement in seven subsequent missions (over a period of 15 years beginning in 2022) would be modest at 250–300 metric tons to low-Earth orbit per mission. Thus two cargo launches of NASA׳s Space Launch System and one crew launch of the Falcon Heavy, for example, would be adequate to maintain support for each consecutive mission. Although we propose developing two sets of electric propulsion systems to account for the Earth-spiral phases, only one vehicle is flown on a heliocentric trajectory at any given time. Thus, our low-thrust stopover cycler with zero encounter velocities falls into a category of a “one-vehicle cycler,” which completes the gap in the literature, where we have already seen multiple-vehicle cycler concepts.

Feasibility Assessment of Rapid Earth–Mars Transfers Using High-Power Electric Propulsion

The aim of this paper is to investigate the feasibility of an Earth-to-Mars transfer with reduced transfer time using a high-power electric propulsion system. The study involved a multidisciplinary analysis combining general performance calculations for power-limited systems, an analysis of a nuclear power source that might be available in the future, and a series of mission analyses, including trajectory optimization. The study quantified the importance of the specific mass of the power and propulsion system with regard to the objective of a fast transfer. A very fast Earth-to-Mars transfer in less than six weeks appears to be unrealistic in the medium term because it depends on a hypothetical breakthrough with the nuclear electric power source, requiring a specific mass that would typically be less than . The results obtained define the limits to improvements that could be obtained in the medium term, using power-generation technologies that, while challenging, are frequently considered as reasonably optimistic, for example using a high-temperature Brayton or Rankine conversion cycle. The shortest Earth-to-Mars transfer time that could be expected for missions with a large enough payload mass was found to be about 120 days, compared to 180 days with chemical propulsion.

Plasma Properties in the Plume of a Hall Thruster Cluster

The Hall thruster cluster is an attractive propulsion approach for spacecraft requiring very high-power electric propulsion systems. Plasma density, electron temperature, and plasma potential data collected with a combination of triple langmuir probes and floating emissive probes in the plume of a low-power, four-engine Hall thruster cluster are presented. Simple analytical formulas are introduced that allow these quantities to be predicted downstream of a cluster based solely on the known plume properties of a single thruster. Nomenclature A = area of one electrode AS = surface area of sheath surrounding an electrode B = magnetic field strength E = electric field strength e = electron charge kb = Boltzmann’s constant me = electron mass mi = ion mass n = electron number density n0 = reference density Te = electron temperature Te,0 = reference electron temperature Vd2 =v oltage measured between triple probe electrodes 1 and 2 Vd3 =v oltage applied between triple probe electrodes 1 and 3 V f = floating potential γ = ratio of specific heats δ = sheath thickness λD = electron Debye length φ = plasma potential φT = thermalized potential Subscript j = contribution from an individual thruster

POWOW — an alternative power source for Mars exploration

Electric propulsion has emerged as a cost-effective solution to a wide range of satellite applications. Robust development of the near-Earth infrastructure that will enable Mars exploration can also rely upon solar-electric powered vehicles. Some of these options will be discussed as well. The POWOW concept is a solar-electric propelled spacecraft capable of significant cargo and short trip times for traveling to Mars. When used for Mars missions, it would enter areosynchronous orbit (Mars GEO equivalent) and beam power to surface installations via lasers. The concept has been developed with industrial partner expertise in high efficiency solar cells, advanced concentrator modules, innovative arrays, and high power electric propulsion systems. The previous spacecraft design providing 898 kW using technologies expected to be available in 2003 produced areal power densities approaching 350 W/m 2 at 80 °C operating temperatures and wing level specific powers of over 350 W/kg were projected. Because of the high power in this satellite, high voltage operation (up to 1000 V) new work on the effects of hypervelocity impact on test modules operated at these voltages is included here. Electric propulsion options have been refined and focus only on high power Hall thrusters of new as well as conventional designs.

Radiometric Analysis from the 26-Kilowatt Electric Propulsion Space Experiment Flight

The U.S. Air Force Research Laboratory's Electric Propulsion Space Experiment (ESEX) was launched and operated in early 1999 to demonstrate the compatibility and readiness of a 30-kW class ammonia arcjet for satellite propulsion applications. As part of this flight, an array of onboard contamination sensors was used to assess the effect of the arcjet and other environments on the spacecraft. The sensors consisted of microbalances to measure material deposition, radiometers to assess material degradation due to thermal radiation or contamination, and solar cell segments to investigate solar array degradation. Over eight firings of the ESEX arcjet, and 33 min, 26 s operating time, the radiometer near the thruster, viewing the arcjet plume and body, experiences a change in the thermal properties of its coating from the firings. Radiometers with no view of the arcjet, or a view of only the plume, show no change. In general, degradation effects are observed only on sensors near the thruster exhaust nozzle, a location unlikely to be used in an operational high-power electric propulsion system. No degradation effects are observed in the backplane of the thruster. For future programs, although engineering measures may be needed for spacecraft equipment in the immediate vicinity of the thruster body, the arcjet environment is generally benign.

POWOW: power without wires: a SEP concept for space exploration

Electric propulsion has emerged as a cost-effective solution to a wide range of satellite applications. Deep Space 1 successfully demonstrated electric propulsion as the primary propulsion source for a satellite. The POWOW concept is a solar-electric propelled spacecraft capable of significant cargo and short trip times for traveling to Mars. It would enter aerosynchronous orbit and from there, beam power to surface installations via lasers. The concept has been developed with industrial partner expertise in high efficiency solar cells, advanced concentrator modules, innovative arrays, and high power electric propulsion systems. The latest version of the spacecraft, the technologies used, and trip times to Mars are presented. The POWOW spacecraft is a general purpose solar electric propulsion system that uses new technologies that are directly applicable to commercial and government spacecraft with power levels ranging from a LEO power level of 4 kW up to GEO spacecraft about 1 MW. The system is modular, expandable, and amenable to learning curve cost projection methods.