Historical highlights

Predecessors of electric propulsion theory go as far back as early 20s century when the American rocket scientist Dr. Robert Goddard first described the possibility of accelerating electrically charged particles to very high velocities without the need for high temperatures. In 1955 Dr. Stuhlinger showed how the lighter-weight electric propulsion systems could make planetary voyages more feasibly than with chemical propulsion. Extensive testing demonstrated that the ion drive engine used to maneuver spacecraft can indeed be used to propel future scientific missions without contaminating the scientific payload. However, these tests were unconfirmed in actual space flight prior to Soviet space program that brought these theories into practice. Soviets first developed xenon ion engines for deep space missions. GSO scientific team members were among the scientists and engineers who designed the group of electric propulsion thrusters used for precision pointing and precision positioning functions. The principles and engineering are now recognized as being applicable to all types of spacecrafts.

Electric propulsion thrusters

Technically, electric propulsion underlines two types of thrusters: Electro-thermal and Plasma thrusters. In turn, Electro-thermal thrusters are subdivided in two categories: Arcjet thrusters and Thermocatalitic thrusters while Plasma thrusters are subdivided in four categories: Stationary plasma thrusters (SPT), Hall thrusters, Pulsed plasma thrusters and Ion thrusters. Here is a brief description of each category.

Arcjet thrusters

Arcjet thrusters are based on a concept of gas acceleration in a supersonic nozzle. Under electrical discharge the heat transfer in a supersonic flow results in significant displacement of the critical throat comparing with classic adiabatic flow. Key parameters like an expansion angle and throat position represent basic functions or arc parameters.

Thermocatalytic thrusters

Thermocatalytic thrusters are smallest in dimension and have incomparable flight proven reliability. Therefore, these are widely used as actuators in spacecraft attitude control systems. Thermocatalytic thrusters consist of a chamber for hydrazine decomposition with catalytic bed, a nozzle, an electric heater and multiple-operating feeding valve.

SPT thrusters

Stationary plasma thrusters are second to thermocatalytic thruster in flight proven reliability. Like thermocatalytic, SPT thrusters were developed and qualified at the Design Bureau FAKEL located in Kaliningrad, Russia during the 1980s. Currently there are three models of flight proven plasma thrusters: SPT-50, SPT-70 and SPT-100 (numerical indexes designate the diameter of accelerating channel in mm). A newer model, the 4.5 kW SPT-140 thruster, was recently added to the SPT family. SPT thrusters are now accepted as the standard for geo-stationary satellites.

Buran/Energia

RSC Energia became interested in electric propulsion studies in 1989 when developmental studies of the universal space platform with on-orbit mass of about 15,000 kg began. It planned to use 22 SPT-100-type thrusters for platform station-stabilizing and attitude control. GSO's team leader, Dr. Vasiliy Starostovich, acquired extensive hands-on experience with RSC Energia/Buran thruster design. Although the project was cancelled, the results of those design studies were used in development of a combined propulsion subsystem for the Yamal-100 geo-stationary and subsequent satellites.

Mars mission

SPT-100 thrusters are scheduled for use in reaching Phobos, with landing, soil sampling and delivery. The interplanetary mission Phobos-Soil, included in the Russia's space program, is supposed to be launched in astronomical window of the year 2005 and completed in 2008. Application of the SPT-based electric propulsion will allow using a light launch vehicle SOYUZ instead of the more expensive PROTON and, furthermore, will save about 100 kg of lander's mass. Different aspects of SPT are under studies at various premier world space research companies. SPT-100s are currently used as baseline design on multiple Western programs.

Hall thrusters

The low-power Hall thruster KM-32 tested for power range from 150 to 400 W and under discharge voltage up to 400 V. The test revealed the specific impulse of 2000s under high-test voltage with thrust efficiency about 45%. Ion energy spectra for the flight prototype of Hall thruster D-55 were experimentally measured with the use of a retarding potential analyzer. The measurement show that Xe ionization occurs within a very narrow layer of voltage variation, for less than acceleration voltage, and that results in very high acceleration efficiency. A Hall-effect thruster ROS-99 is currently undergoing an accelerated lifetime test in Russia.

Pulsed plasma thrusters

Pulsed plasma thrusters (PPT), or more precisely, the ablative pulsed plasma thrusters have been well known since the mid-60s. Intensive studies for use of these thrusters onboard for small satellites, i.e. satellites with total mass of 25-250 kg and the power available for propulsion system of 10-150W, were performed. Propulsion systems, based upon the pulsing plasma thrusters, could be even more mass efficient than widely used SPT-based systems for the total thrust impulse ranging from 50 up to 200 kNs. For a small satellite of total mass of 500 to 1000 kg, these thrusters could be used for final injection into the operating orbit and deorbiting, GEO stationkeeping, or LEO aerodynamic drag compensation, and for high-accuracy attitude control. Experimental results that could improve the understanding of plasma formation and acceleration were obtained in series of experiments with laboratory models of PPT for the pulse energy range from 20 up to 1000J. The results were correlated with the one-dimensional magneto-hydrodynamics model (MHD) of accelerated plasma. It was shown that the improved conditions for accelerated plasma were capable of nearly doubling the thrust efficiency. This area is within the most progressive research interests of many teams including us at GSO.

Ion thrusters

Russian researchers presented results of numerical simulation of a low-power Xe-ion thruster with an advanced, slit-type accelerating system. The analysis were verified on a 5sm laboratory model of the ion thruster, equipped by a system of two grids made from 0.5mm molybdenum and accelerator made from pyrolitic graphite with a thickness of 1mm. Experiments were carried out for the power range of 50-150 W and specific impulse values of 2500-3500s were achieved. The values of thruster efficiency reached nearly 60%.

Problems of spacecraft/thruster interactions

More advanced thruster's designs increase precision attitude control and precision orbit adjustment in space, they have shown themselves to be more weight and cost effective, and promise increased propellant efficiency. However, it is still very important to conduct extensive supplementary testing on new thruster models' interactions with other parts and equipment housing spacecraft to assure overall safety. Some experimental studies have shown that advanced thrusters' operation may adversely interact with spacecraft safety. For instance, researchers from Keldysh Institute of Applied Mathematics presented the results of a numerical simulation of plume interactions with micro satellites. Due to a non-traditionally proximate placement of the attitude control hydrazine thrusters, there was a threat of contamination of the satellite's sensitive surfaces by plume products. Numerical results derived from quasi-gasdynamic model along with results of a parallel experimental study confirm the absence of the re-circulation zone and thus contamination free safety of the accepted design.

A traditional problem for satellites with wide-spanned solar arrays in the far-field interaction of quasi-neutral plasma plume with the arrays that could result in thrust reduction and spurious torque. The dynamic interaction of SPT-70 plume with the 250 by 450 mm model of solar wing has been experimentally studied by researchers from Design Bureau FAKEL. The proposed technique of predicting dynamic pressure of plasma plume showed satisfactory agreement with the experimental data. Another experimental study showed that products of thruster internal parts' erosion may also contaminate spacecraft surfaces. For the typical geometry of a GEO telecom satellite, erosion of the internal ceramic isolator in SPT-100 may cause deposition contamination films on certain solar cells.

The external surfaces of spacecraft are also subjected to erosion in SPT exhaust plumes. Lifetime study has shown significant erosion of Kaptan film, widely used as an external layer of MLI blanket. Samples were exposed to SPT-70 plume during a 35 hours period.

It here bears noting that ISP (Interaction Spacecraft Propulsion) and ESCAPE (Electro-Static Charge in Artificial Plasma Environment) specialized software packages developed by Russian software development teams are widely used not only in Russia but in every country concerned with space systems development.

Future of Electric Propulsion

Despite numerous definite conclusions made within plasma sources for space propulsion, this area remains underdeveloped in terms of deep space travel. Both theoretical and experimental activities present many opportunities for investigation where the most critical aspects are operation of plasma sources and their compatibilities with various spacecraft systems. Theoretical studies necessitate development of new mathematical tools for further investigation of the physical processes taking place inside the plasma sources and the mechanical and electromagnetic impacts of the plasma sources at system levels. Special efforts are being brought to bear on the theoretical explanation and modeling of the plasma discharge process on a radio frequency plasma source. Further development of possible applications of these plasma sources to the solution of concrete tasks in space are necessary. To this end, new mathematical algorithms must be developed and implemented.

The investigation of the physical processes and characteristics of a medium power hall thruster, a pulse plasma thruster, an arcjet thruster with hollow cathode, and some new high frequency ion sources, as well as their possible application have still numerous gray areas. Existing mathematical models of the plasma dynamics inside the SPT accelerating channel with a kinetic description of the ion, neutral, and electron dynamics, models of the plasma interaction with the discharge chamber walls and the wall erosion, a model of the sputtered particle dynamics inside the SPT accelerating channel and of its plume require ongoing investigation.

While prospects of all the studied electric propulsion thrusters are very definite, newer applications on both theoretical and software level should be pursued to increase accurate momentum control, pointing, vibration isolation, guidance and navigation, data control and other satellite operational properties. Our team at GSO is fully equipped to assist research efforts in the above described areas.