ION THRUSTER OPTICS DESIGN FOR THE
INTERSTELLAR PRECURSOR MISSION
Brittany Albrandt, Donny Newsom
Colorado State University
ABSTRACT
The NASA Origins Program, which seeks an understanding of the origin of the universe, requires deep space (interstellar) exploration. The Interstellar Probe Mission is a part of this program, which will require delivery of a spacecraft to a location 250 astronomical units from the Earth in 10 years. One mission scenario would involve launch around 2007 and subsequent interstellar travel sustained by a Nuclear Electric Propulsion (NEP) package. The NEP scheme involves the production of electrical power from a nuclear heat source and consumption of the power in one or more ion thrusters. This paper will describe ion thruster operation and explain why NEP with ion thrusters yields performance superior to that of conventional chemical thrusters in this application. The selection of krypton propellant and an ion acceleration voltage of 13 kV will be justified. Detailed numerical analyses will be presented that show the trajectories of ions as they are extracted through electrodes (the ion optics) and accelerated to the high exhaust velocities (high specific impulses) at which the thrusters must operate to perform this mission efficiently. Results obtained from these analyses will be used to define the ion optics geometry that will assure proper thruster operation over the full 10-year mission lifetime and spacecraft acceleration to the desired terminal velocity of the probe.
INTERSTELLAR PRECURSOR MISSION
The origins program seeks to answer two enduring questions. Where do we come from? Are we alone? “NASA’s Origins Program has a goal (by ~ 2040) of imaging and remote sensing the potentially Earth-like planets around other stars out to approximately 40 light years.” [i] The program is based on a family of missions. These missions include some ground-based observatories and some spaced-based observatories. The spaced-based observatories are broken up into four groups of missions; Precursor missions, First-generation mission, Second-generation mission, and Third-generation mission. The early missions are designed to determine which planets have the capability of sustaining life. Subsequent missions are designed to study these planets in greater detail. An example of one of the first precursor missions was the Hubble Space Telescope (HST). The Interstellar Probe mission, expected to launch in 2007, is designed to travel to a distance of 250 astronomical units in 10 years. An astronomical unit is defined as the mean distance between the sun and the earth, which is 1.496x1011m. The goals of the Interstellar Probe are to examine the characteristics of some of these planets while developing the technology for later missions.
NUCLEAR ELECTRIC PROPULSION (NEP) VS. CHEMICAL PROPULSION
Some options for propulsion systems include nuclear electric propulsion and chemical rockets. NEP involves an electric power generator powered by a nuclear fission heat source. Thrust is produced by acceleration of ions through the device. An examination of chemical propulsion system versus a NEP system follows. The basis of this analysis is Equation (1).
(1)
Where mi is the initial mass of the rocket, including propellant (kg); mf is the final mass of the system (kg); Vvelocity is the change in spacecraft velocity (m/s) required to accomplish the mission; and Ue is the exit velocity of the propellant (m/s). The law of conservation of energy, expressed as Equation (2), can determine the exit velocity for the NEP propellant:
(2)
Where Vvoltage is the difference in voltage through which a positively charged propellant ion is accelerated through (V); q is the charge of an ion (C); m is the mass of the propellant ion (kg); and Ue is the exit velocity (m/s). Solving the above equation for Ue:
For example, if the discharge chamber is charged to 13,000V and krypton (atomic weight = 83.8) is used for the propellant, the average exit velocity of an ion would be:
Ue 180,000m/s
This result assumes that all of the propellant atoms are ionized and then accelerated to produce thrust. Laboratory work has shown that not all atoms are ionized to contribute to the thrust. The true mean exit velocity, as a consequence, is reduced. In the present case, this means the true exit velocity is expected to be 140,000 m/s. Equation (2) also demonstrates why krypton, and not xenon, is the propellant of choice for this mission. In order to achieve the same exit velocity (180,000m/s) for xenon (atomic weight = 131), the voltage must be larger than 13,000V. Thus, krypton was chosen to obtain lower voltages that present less challenging design issues. It also should be noted that the Vvelocity term is different for the ion thruster and the chemical rocket. One of the main differences between a chemical rocket and an ion thruster is that a chemical rocket thrusts intensely for a short amount of time, where as an ion thruster thrusts continually at low level for the duration of the mission. Figure 1shows a typical velocity distribution for a chemical rocket and ion thruster.
For the chemical rocket, the average Vvelocity can be approximated by equation (3).
(3)
Where Vavg is the average change in velocity (m/s) for a chemical rocket; s is the distance traveled (m); and t is the mission duration (thruster time) (s). For a typical chemical rocket, traveling a distance of 250AU at a constant velocity:
Vchemical = 1.19x105 m/s
Using this value in equation (1) and assuming a final mass of 1,000 kg and an exhaust velocity (for a typical chemical rocket) of 4,000 m/s:
mi = 8x1015 kg
Nearly all of this initial mass is propellant required to produce the needed thrust. This mass is so great that it makes the mission impractical.
The NEP system thrusts throughout the entire mission and must still realize the same average velocity as the chemical system. The Vvelocity value for the NEP becomes:
VNEP = 2Vchemcial (4)
Using the V found for the chemical rocket:
VNEP = 2*1.19x105 m/s
VNEP = 2.38x105 m/s
The initial required mass can then be calculated by assuming a final mass of 1,000 kg payload, a 9,000 kg mass for the nuclear power plant, and an exit velocity of 140,000 m/s, Equation (1) becomes:
mi = 50,000 kg
This comparison clearly shows the benefits of NEP. For a mission to 250AU, chemical propulsion is not a viable option due to the excessive propellant mass required to accelerate the payload. The initial mass for the nuclear electric propulsion is a reasonable amount of mass.
OPERATION OF ION THRUSTERS
Figure 2. Schematic of Typical Ion Thruster
The operation of an ion thruster is a fairly simple process that involves simple electrostatic forces. Propellant, typically xenon or krypton, enters the ionization chamber. Inside the chamber, the propellant is bombarded by energetic particles to form positively charged ions. Both the screen grid and the acceleration grid are made up of thousands of small, round holes. As the positive ions are accelerated through the grids by the potential difference between the two electrodes, a thrust is produced. The ionization chamber is typically charged to some large positive voltage. For the Interstellar Probe Mission, the chamber and the screen grid are charged to approximately 13,000V. The positively charged ions, or discharge plasma, are then accelerated out of the chamber and through the screen grid and the acceleration grid. Negatively charged electrons produced from each ionization event are collected at the anode and removed via a power supply to the neutralizer. From there they are emitted into space at the same rate as the ions being produced. The acceleration grid is sufficiently negative compared to the neutralizer and discharge plasma to prevent electrons from being drawn upstream through the holes into the discharge plasma.
A key performance indicator of a propulsion system is its specific impulse, or thrust produced per unit of propellant consumption. The specific impulse is determined from the mean exhaust velocity of the propellant (Ue) by the expression (5):
(5)
Where g is the acceleration due to gravity (m/s2). For the case of the ion thruster where the exhaust velocity is 140,000m/s then:
Isp 14,000s
The higher the specific impulse of the ion thruster compared to that of the chemical rocket (400s) indicates its more efficient performance.
GRID DESIGN
Figure 3. Grid Schematic
Figure 3 shows a cross-section for one of the hole sets in grids. The upstream screen grid, as well as the discharge chamber and plasma is charged to a potential of 13,000V. The accelerator grid is charged to –500V. The diameter or the screen grid holes (ds), diameter of the acceleration grid hole (da), thickness of the screen grid (ts), thickness of the accelerator grid (ta), and the distance between grids (lg) must all be analyzed in order to design an efficient thruster system.
DESIGN ISSUES
A major design requirement is that the ion thruster grids must have a lifetime of more than 10 years. The grids erode rapidly when ions strike them and sputter away the grid materials. Figure 4 illustrates a one way in which this can occur.
Figure 4. Direct Impingement
Figure 4 shows a schematic of direct impingement. This phenomenon occurs when too much current is forced through the acceleration grid holes. The schematic shows the upper half on one grid hole. The lines represent trajectories of individual ions as they traverse the length of the thruster. As the figure shows, ions collide and will erode the grids over long mission durations. As the grid material erodes, small bits of material may become dislodged and create a short between the grids. A second factor influencing the lifespan of the grid is cross-over erosion, which is also a form of direct impingement. Figure 5 shows this schematic:
Figure 5. Cross-Over Erosion
In this scenario, the ions impinging the accelerator grid originate from other half of the sheath. This can also cause grid erosion. Figure 6 shows an optimized ion thruster without any form of impingement.
Figure 6. Ion Thruster
Another design criterion is the minimum voltage at the centerline of the grid must be –30V. This negative voltage insures that the electrons from the neutralizer and downstream plasma do not turn around and reenter the plasma chamber. Figure 7 shows these design criteria graphically.
Figure 7. Voltage Criteria
Where Vn is 13,000V and Vt is 500V. The minimum, labeled as “Potential Through Apertures,” represents the design criteria of –30V.
Nuclear electric propulsion, although more effective than chemical propulsion for interstellar missions, does have some political drawbacks. The ion thruster system is powered by a nuclear reactor. Although the Soviet/Russian space programs have been using nuclear reactors is space exploration for quite some time, nuclear power carries a stigmatism. Some people feel that nuclear power is dangerous not only to human beings but also to the environment, both here and in space. Also, nuclear power is strictly regulated in the United States. In order for a nuclear reactor to be launched into space by the United States, the president must sign off on the nuclear reactor prior to launch.
CONCLUSION
Considering all the geometry and design criteria an optimal design can be selected. For this case, a potential of 13,000 volts for the plasma, accelerator grid voltage of –256 volts, and a centerline voltage of –30, the optimal geometry is:
Grid Spacing 9 mm
Screen Hole Diameter 9mm
Acceleration Grid Hole 4mm
Acceleration Grid Thickness 4mm
Screen Grid Thickness 1.5mm
Hole-to-Hole Distance 9.5mm
Normalized Perviance per Hole 1-1.5
These results are based on laboratory tests performed in another undergraduate research study. The designs tested were direct results of the analysis discussed in this paper.
Nuclear electric propulsion is the future of space exploration. It provides a more effective propulsion system than traditional chemical systems. Careful analysis of the ion optics leads to an optimized design for the Interstellar Probe Mission.
1
[i] Patterson, M.J., R.F. Roman, and J.E. Foster. “AIAA 2000-3811: Ion Engine Development for Interstellar Precursor Missions.” NASA Glenn Research Center: July 16-19, 2000.