Electric Propulsion Options Research Articles

Solar Electric Propulsion for Discovery-Class Missions

This paper offers a user-centric consolidation and comparison of the full range of government and commercial solar electric propulsion options available in the near term for primary propulsion on deep-space science missions of the class commonly proposed to NASA’s Discovery program. Unlike previous papers, this work does not emphasize feasibility from a mission-analysis perspective. Rather, it emphasizes requirements uniquely imposed by competitively reviewed cost-capped mission proposals, for which system-level flight heritage and cost credibility can trump sheer performance and mission capture. It describes criteria that mission architects and review boards can use to select and evaluate electric propulsion systems, provides descriptions of viable government and commercial electric propulsion system options, describes the modifications needed to adapt commercial electric propulsion systems to deep space, and discusses appropriate methods for costing commercial-based electric propulsion systems. It concludes that the solar electric propulsion systems best suited for Discovery missions have a solid system flight heritage that can meet the requirements for deep space with minimal modifications. Commercially developed electric propulsion systems offer significant heritage potential and, in many cases, the required changes for deep-space application introduce comparatively low technical and cost risk.

MHD instabilities in magneto-plasma-dynamic thrusters

Magneto-plasma-dynamic thrusters (MPDTs) act as electromagnetic plasma accelerators and represent a high power, electric propulsion option for primary space missions. One of the major problems facing MPDT operation is the onset of a critical regime, which is found when the power is increased beyond a threshold value, depending mainly on the thruster geometry, the type and mass flow rate of the propellant and the intensity of the magnetic field applied. In this regime, large fluctuations in the electrode voltage signals, damage to the anode and decreased efficiency are observed. Since 2000, several test campaigns have been carried out to investigate the electrostatic and magnetic properties of plasma fluctuations using electromagnetic and optic probes and ultraviolet tomography. Results obtained have shown a strong relation between the onset phenomena and the growth of a large scale magnetohydrodynamic (MHD) instability, with helical kink mode features. On the basis of the experimental observations, a passive method to suppress instability is proposed and has been partially tested, with encouraging results.

Geosynchronous-Earth-Orbit Communication Satellite Deliveries with Integrated Electric Propulsion

DOI: 10.2514/1.35322 The use of electric propulsion was evaluated for transfer of communication satellites from geosynchronous transfer orbits to geosynchronous earth orbits. Recent communication satellite designs, normal launch vehicle delivery orbits, and integrated electric propulsion subsystems (with input powers less than nominal satellite power levels) were assumed to minimize required changes to present and near-term launchers and spacecraft. The capture fraction of recent communication satellites that could have been delivered was evaluated versus launcher delivery capability, launch site, and in-space propulsion characteristics. Electric propulsion significantly increases the capturefractionoflauncherswithgeosynchronoustransfermassdeliverycapabilitieslessthan(dependentonlaunch site) about 4500 to 5500 kg. Insertion times at given launch sites were found to be accurately specified by the satellite power-to-mass ratiosandtheassumedelectricpropulsionspecific-impulse/efficiencycharacteristics. Insertiontimes less than 100 days were found for satellites with high power-to-mass ratios that used high thrust-to-power electric propulsion options. The influence of power for electric propulsion beyond that used for satellite payloads and housekeeping was also assessed, and insertion times less than a month will require powers significantly higher than presently installed.

On the design of low‐thrust transfers through L1 for the ESMO mission

AbstractThe following paper summarises the mission analysis studies performed for the phase A of the solar electric propulsion option of the European Student Moon Orbiter (ESMO) mission. ESMO is scheduled to be launched in 2011, as an auxiliary payload on board an Ariane 5. Hence the lack of influence on the launch date. Two methods were devised to assess efficiently wide launch windows for the Earth‐Moon transfer, they are presented below. (© 2008 WILEY‐VCH Verlag GmbH & Co. KGaA, Weinheim)

Alternate Direct-Drive Option for Electric Propulsion

The goal of this research is to power a 1.5-kW Hall thruster with a power source based on chemical double layer capacitor technology and to access performance characteristics of this power/propulsion system. The thruster was operatedinalong-pulsemodewithdecayingpowerinputandexhibitedsimilarperformancecharacteristicstobeing operatedundersteady-stateconditions.Overthetestedpowerrangeof450–1750W,thethrustergeneratedthrustof 30–100 mN at an efficiency of 30–55% and a specific impulse of 900–1950 s. These results formed the basis for a mission analysis considering extremes with regard to shading and atmospheric density. A near-Earth orbiting spacecraft served as an analysis platform to compare the mission performance of direct-drive electric propulsion options and a state-of-the-art hydrazine monopropellant system. Regardless of orbit conditions, electric propulsion options outperformed the conventional chemical system in terms of wet mass decrease (41–59%) and net mass increase (14–47%), while capacitor-powered electric propulsion options substantially reduced the onboard power requirement (66–91%).

Design of Earth–Mars transfer trajectories using evolutionary-branching technique

In this paper a novel global optimisation approach is used to search for potentially interesting solutions for a mission to Mars. The approach blends the characteristics of evolutionary algorithms with the systematic search, typical of branching techniques. Solutions for a roundtrip to Mars, either direct or via Venus, considering long and short stays on Mars or free return trajectories are analysed, thus providing a comprehensive view of all the opportunities within a wide range of possible launch dates. Finally, electric propulsion options are investigated including the possibility of using Mars’ Lagrangian points for a low cost capture. The proposed global search was effectively able to find globally minimal Δ v solutions for Earth–Mars roundtrips giving the expected characterisation of the problem under study. Moreover, it will be shown how the method was able to automatically rediscover known solutions along with new ones of practical interest for future Mars exploration.

A survey of propulsion options for cargo and piloted missions to Mars.

In this paper, high-power electric propulsion options are surveyed in the context of cargo and piloted missions to Mars. A low-thrust trajectory optimization program (raptor) is utilized to analyze this mission. Candidate thrusters are chosen based upon demonstrated performance in the laboratory. Hall, self-field magnetoplasmadynamic (MPDT), self-field lithium Lorentz force accelerator (LiLFA), arcjet, and applied-field LiLFA systems are considered for this mission. In this first phase of the study, all thrusters are assumed to operate at a single power level (regardless of the efficiency-power curve), and the thruster specific mass and power plant specific mass are taken to be the same for all systems. Under these assumptions, for a 7.5 MW, 60 mT payload, piloted mission, the self-field LiLFA results in the shortest trip time (340 days) with a reasonable propellant mass fraction of 57% (129 mT). For a 150 kW, 9 mT payload, cargo mission, both the applied-field LiLFA and the Hall thruster seem reasonable choices with propellant mass fractions of 42 to 45%(7 to 8 mT). The Hall thrusters provide better trip times (530-570 days) compared to the applied-field LiLFA (710 days) for the relatively less demanding mission.

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.

Chemical and solar electric propulsion options for a cornerstone mission to mercury

In its recent meeting in Naples, ESA has confirmed a mission to Mercury as the conrerstone element of its Horizon 2000+ program for solar system science. The main trade-offs of the mission analysis are presented. 8 missions opportunities with chemical propulsion have been found between 2006 and 2012, with cruise durations of 6 to 7.8 years. The selected option provides much shorter cruise times of 2.1 to 2.6 years. It combines gravity assists with solar electric propulsion (SEP) for the cruise phase and chemical propulsion for insertion and orbital manoeuvers in orbit around Mercury. This strategy which provides the best compromise between the mass budget and the cruise time will be tested at the end of 2002 by the SMART-1 mission recently approved by ESA, which combines SEP with lunar gravity assists.

Propulsive options for a manned Mars transportation system

In this investigation, five potential manned Mars transportation systems are compared. These options include: (1) a single vehicle, chemically propelled (CHEM) option, (2) a single vehicle, nuclear thermal propulsion (NTP) option, (3) a single vehicle solar electric propulsion (SEP) option, (4) a single vehicle hybrid nuclear electric propulsion (NEP)/CHEM option, and (5) a dual vehicle option (NEP cargo spacecraft and CHEM manned vehicle). In addition to utilizing the initial vehicle weight in low-earth orbit as a measure of mission feasibility, this study addresses the major technological barriers each propulsive scenario must surpass. It is shown that instead of a single clearly superior propulsion system, each means of propulsion may be favored depending upon the specified program policy and the extent of the desired manned flight time. Furthermore, the effect which aerobraking and multiple transfer cycles have upon mission feasibility is considered.