Electrothermal Propulsion Systems
Electrothermal thrusters differ from both electromagnetic and electrostatic propulsion systems due to their operational design; electromagnetic and electrostatic systems propel charged ions through the use of electric and magnetic fields, while electrothermal systems heat the propellant, and rely upon thermal dynamics to propel the system (Jordan, 2000). In typical operation, a propellant is electrically heated, which increases the pressure and expands the gas, forcing the energized mass out of the nozzle and providing thrust to the spacecraft (Jahn & Choueiri, 2002). There are three types of electrothermal propulsion systems; arcjet, resistojet, and inductively or radiatively heated systems (European Space Agency, 2004). Each of these three propulsion systems is considered viable, and each is discussed in the succeeding paragraphs.
Resistojet Thruster
Figure 10: A cross section of a resistojet thruster (Jordan, 2000).
A resistojet propulsion system relies upon the outlined concept for electrothermal thrusters; as shown in Figure 10, a propellant is injected into the resistojet assembly, heated by a power supply and radiative heat transfer, and then expelled through a compressive exhaust nozzle (Jordan, 2000). In more advanced resistojet concepts, a dual stage reaction increases propellant efficiency; hydrazine, after undergoing conventional chemical decompression, is then electrically heated and expelled, massively increasing the thrust yield from a set amount of propellant (European Space Agency, 2004; Jordan, 2000). However, the Isp for resistojet thrusters is limited due to the high molecular mass of the propellant and an approximate 3000 K temperature ceiling due to direct exposure of critical elements to internal resistojet temperatures (European Space Agency, 2004). As noted in Table 1, an electrotheremal resistojet has an approximate 150 to 700 Isp range, efficiency levels between 35% and 90%, and thrust between 5 and 5,000 mN; for an resistojet powered by a dual stage hydrazine reaction, Isp becomes restricted to between 299 and 304, energy conversion efficiency exceeds 300% (due to additional thrust provided by the decompression of the hydrazine), and thrust levels of 330 mN and above (Jordan, 2000). Both the high efficiency levels and high maximum thrust levels indicate potential viability as described in this report, but a low Isp and negligible systems development or testing within the report parameters detract from the overall operability of the resistojet propulsion system. Nevertheless, research by the Georgia Tech College of Engineering indicated the resistojet’s viability for a high mass Martian transit mission (Seitzman, 2006), further indicating the viability of this technology within the primary constraints.
Arcjet Thruster
Figure 11: Cross section diagram showing the operation of an arcjet thruster (Jordan, 2000).
Another potentially viable electrothermal propulsion system is the arcjet thruster. This propulsion system relies upon a centrally located cathode surrounded by an anode; a high voltage electric field is generated between the cathode and anode, while the propellant is injected between the two electrodes (Jordan, 2000). As shown in Figure 11, the anode additionally functions as the nozzle for expelling the superheated gas (Jordan, 2000). In the resistojet propulsion concept, a primary limiting factor was the maximum attainable operating temperature of 3,000 K; this temperature maximum is increased to between 10,000 K and 20,000 K in arcjet thrusters (European Space Agency, 2004). The indirect application of these extremely high temperatures through an electric current provides insulation for operational elements, and the expulsion pattern of the superheated gas contributes to the insulating effect (Jahn & Choueiri, 2002). Comparatively large Isp ranges, between 280 and 2,300, give significant advantages over resistojet propulsion methods, while a comparable thrust range of 50 to 5,000 mN further indicates the viability of the arcjet propulsion system (Jordan, 2000). Conversely however, the resistojet thruster system provides a significantly higher energy conversion efficiency range, of approximately 35% to 90% compared with the arcjet efficiency range of 30% to 50% (Jordan, 2000). Additional concerns detract from the operability and viability of the arcjet system, including electrode erosion and massive power requirements, has led to negligible implementation or development of the thruster in relation to use as a primary propulsion method (Jahn & Choueiri, 2002). While this report considers the arcjet thruster as a more viable alternative to the resistojet system, lacking implementation and moderate performance reduce its ability to meet the outlined requirements adequately.
Inductively or Radiatively Heated Thruster
Figure 17: A cross section showing the operation of a VASIMR propulsion system (NASA, 2003).
The final method for electrothermal propulsion is inductively or radiatively heated systems. As a major limiting factor for arcjet implementation is continual reliability due to electrode corrosion, inductively or radiatively heated systems utilize an applied, oscillating electromagnetic field to interact with the propellant and generate primary thrust (Jahn & Choueiri, 2002). Propulsion systems within this category include both electrothermal and electrostatic components, and some thruster types include additional electromagnetic propulsion aspects; for this reason, classification of the subsequent systems is often inconsistent (Jahn & Choueiri, 2002). Several prototypes for this propulsion type are currently being developed, including the Variable Specific Impulse Magnetoplasma Rocket (VASIMR). The VASIMR system produces extremely high Isp levels, theoretically between 3,000 and 30,000 and tested between 5,000 and 12,000, relatively high tested thrust levels of 5,000 mN, and moderately high efficiency levels of approximately 60% (Ad Astra Rocket Company, n.d.; Bering, et al., 2008; Carter, et al., 2005; Jordan, 2000). This system includes electrothermal and electromagnetic propulsion components, indicating its placement within the inductive and radiatively heated systems category. The operation of the VASIMR system relies upon the ionization, compression, acceleration, and expulsion of a propellant gas (Ad Astra Rocket Company, n.d.). A high energy radiofrequency, or helicon, coupler ionizes the upstream injected propellant gas, creating positively charged plasma, which is then accelerated and compressed to a secondary radiofrequency chamber (Ad Astra Rocket Company, n.d.). This secondary radiofrequency coupler, part of the Ion Cyclotron Heating (ICH) section, energizes the compressed plasma, and increases the temperature of the plasma to approximately 1,000,000 K (Ad Astra Rocket Company, n.d.; Bering, et al., 2008). The highly energized plasma is then directed into the magnetic nozzle, where ionic orbital momentum is converted into linear momentum and expelled from the thruster; exhaust velocities can reach approximately 50 kilometers per second, and generating considerable amounts of thrust; a visual representation of the operation of the VASIMR system is depicted in Figure 12 (Ad Astra Rocket Company, n.d.). In addition to high performance capabilities, the VASIMR propulsion system conditionally varies the power levels of the two radiofrequency couplers, increasing or decreasing both Isp and thrust, depending upon mission constraints (Ad Astra Rocket Company, n.d.; Carter, et al., 2005). Several principal concerns reduce the viability of the VASIMR system, including the safety concern of crewmember exposure to intense electromagnetic and radiofrequency radiation, and the operational concern considering the currently inadequate stage of development and implementation; space-based testing has not yet been achieved with this system (Jordan, 2000). In summation, the VASIMR propulsion system provides overall increases in performance over comparative systems while allowing significant in-flight variability, but space-based testing and further development is required before the system is fully operational.
Electrothermal Propulsion Systems: Conclusion
Of the various electrothermal systems discussed, the performance of the VASIMR system provides unparalleled advantages over both resistojet and arcjet systems. While comparable thrust is theoretically achieved by all three systems, VASIMR remains the only tested technology to produce 5,000 mN of thrust; additionally, testing places VASIMR Isp levels significantly higher than both other systems, between 5,000 and 12,000 compared to either 280 to 2,300 or 150 to 700 (Jordan, 2000). The extremely high Isp levels prove highly advantageous for long distance transits, while the relatively high thrust levels indicate moderately fast acceleration. While outstripping both comparative systems in performance, the VASIMR system requires significantly more development before full operability is achieved; equivalently, the arcjet and resistojet systems are not developed or tested at significantly high performance levels. Therefore, this paper recommends the VASIMR technology as the most viable electrothermal propulsion system for meeting the mission constraints, even though the technology remains to be fully developed. As both the resistojet and arcjet systems are negligibly developed for the mission constraints, they are not considered fully viable by this report, and so are excluded from recommendation.