Linear gridless ion thruster

Details Linear gridless ion thruster Linear gridless ion thruster

2002-06-13

2003-11-04

Kim, Ted (Department: 3746)

Power plants

Reaction motor

Ion motor

C313S362100, C315S111610

Reexamination Certificate

active

06640535

ABSTRACT:

FIELD OF THE INVENTION
The present invention relates to propulsion systems and, more particularly, to a linear gridless ion thruster, which combines an ionization stage from a gridded ion thruster and an acceleration stage from a closed-drift Hall thruster to take advantage of the strength of both thrusters without suffering from the weakness of either.
BACKGROUND OF THE INVENTION
The Rocket Equation (Equation 1):
Mf
Mo
=
Exp
⁡
(
-
Δ
⁢
 
⁢
V
gIsp
)
Equation
⁢
 
⁢
1
shows that the ratio of payload or final mass (Mf) over initial mass (Mo) depends on the velocity increment (&Dgr;V) needed for a spacecraft, and the speed at which exhaust propellant leaves the propulsion system; also known as specific impulse (Isp), which is proportional to propellant exhaust velocity through the gravitational constant (g). That is, the amount of propellant needed to achieve this &Dgr;V is reduced if the Isp of the propulsion system is increased. For example, cryogenic chemical rocket motors such as the Space Shuttle Main Engine are capable of producing specific impulses of about 450 seconds. Chemical rockets employed for long-duration space voyages must use non-cryogenic propellants that yield lower performance (<330 seconds).
Studies have shown that ideally, an engine that would be used as the primary source of propulsion for orbit transfer missions or for satellite station-keeping should produce an Isp between 1000 and 2000 seconds. Spacecraft propulsion systems for interplanetary missions may need to generate even higher exhaust velocities. To achieve the desired performance, a propulsion system must accelerate a propellant gas without relying on energy addition through chemical reactions.
One approach is the application of electrical energy to a gas stream in the form of electrical heating and/or electric and magnetic body forces. This type of propulsion is commonly known as electric propulsion (EP). EP can be categorized into three groups. Electrothermal Propulsion Systems electrically heat a gas, either with resistive elements or through the use of an electric arc, which is subsequently expanded through a nozzle to produce thrust. Electromagnetic Propulsion Systems use electromagnetic body forces to accelerate a highly ionized plasma. Electrostatic Propulsion Systems use electrostatic forces to accelerate ions. In addition to possessing suitable exhaust velocities, an EP system must be able to convert onboard spacecraft power to the directed kinetic power of the exhaust stream efficiently.
To show the benefit of EP systems over chemical systems reference is made to FIG.
1
.
FIG. 1
is a plot of the Rocket Equation showing the final-to-initial mass ratio for a number of missions that use conventional propulsion systems. Clearly the smaller the mass ratio, the more expensive a mission becomes. While missions to Low Earth Orbit (LEO), the moon, and Mars require significantly more propellant mass than payload mass when using chemical propulsion systems, this is not the case for EP systems due to their high Isp. This fact translates into significant cost savings for commercial, military, and scientific space missions.
FIG. 2
shows payload mass and fraction delivered to Geosynchronous Earth Orbit (GEO) as a function of trip time for EP and chemical propulsion systems assuming a moderate launch vehicle (Atlas IIAS) is used.
FIG. 2
compares the performance given by a bi-propellant chemical rocket (Isp=328 sec), an arcjet using hydrazine decomposition propellant (Isp=600 sec), and a Hall thruster using xenon propellant (Isp=1600 sec). As
FIG. 2
clearly shows, the amount of payload delivered to GEO increases with Isp and with trip time. The former is because the launch vehicle places a fixed spacecraft mass in LEO and as Isp increases, the amount of propellant needed for the transfer reduces. The mass that was used for propellant in the all-chemical spacecraft can now be used for payload.
A 15% increase in payload mass can be realized by simply using EP for North-South stationkeeping (NSSK) and using chemical propulsion for the LEO-to-GEO transfer. While the LEO-to-GEO trip takes longer with more of the transfer being done with EP, less propellant is required. Hence, the high-Isp EP system is used more for longer transfers, and more payload can be delivered to GEO.
This principle is being considered for the human exploration of Mars. NASA has now expressed an interest in developing the capability to send a crew to Mars within the next two decades. However, mission cost is a clear driver. Since the LEO-to-MTO (Mars Transfer Orbit) &Dgr;V is a significant fraction of the total mission &Dgr;V, and hence accounts for much of LEO initial vehicle mass, NASA has baselined the use of a Solar Electric Propulsion (SEP) stage to raise a chemically-powered Mars Transfer (MT) stage to a highly elliptic orbit around the Earth. Once the MT stage is in the proper orbit, the crew uses a small, chemically-propelled vehicle to rendezvous with it. Once the crew is in place and the MT stage has been certified to be fully operational, it separates from the SEP stage and ignites its engines for the trip to Mars.
EP's resurgence in recent years is due both to the public's interest in space exploration and money that be saved by commercial spacecraft developers. As illustrated above, the latter comes by virtue of the fact that EP's large specific impulse means that it can accomplish a mission with less propellant than conventional propulsion systems. The recent successes of the Deep Space-1 and Mars Pathfinder missions have helped to renew the public's excitement about space exploration.
The Mars mission scenario described above reduces both trip time (for the crew) and initial spacecraft mass by utilizing a high-performance SEP stage. The key to developing the SEP stage is the utilization of an engine that posses high specific impulse, high thrust efficiency, and a large range of specific impulse over which it can operate while maintaining high efficiency.
At first glance, a gridded ion engine appears to be ideal for the above application. Ion thrusters have very high specific impulses and efficiencies, and have a moderately large range of specific impulses over which they can operate at better than 50% efficiency. However, since such an engine will need to process hundreds of thousands or millions of watts of power, conventional gridded-ion thrusters are inappropriate given the size requirement such an engine would have due to its space-charge and grid erosion limitations.
On the other hand, conventional single-stage Hall thrusters possess high engine efficiency at moderately-high specific impulses. However, the ability to operate single-stage Hall thrusters with long life at very high specific impulses has never been demonstrated nor can ionization processes be decoupled from acceleration processes. The latter results in the strong interdependence of discharge current, discharge voltage, and propellant flow rate that limits the operational flexibility of these engines.
Furthermore, since ions are created at various spots along the ionization/acceleration region, not all ions benefit from the full accelerating potential of the discharge, resulting in a loss of engine efficiency. Moreover, the effect on engine life of placing 1000-2000 V discharge voltages on single stage Hall thrusters (e.g., on the anode from back-streaming electrons) is unknown. Lastly, for specific impulses of ~1300 seconds or less, conventional Hall thruster efficiencies are low because of the coupled ionization and acceleration zones. This would serve to limit the “throttling” capability of the SEP stage (e.g., to provide “high” thrust at moderate specific impulse for certain phases of its orbital burn).
The desire for high throttling performance (also known as “Dual Mode Operation”) applies to a number of commercial, military, and scientific missions. For commercial and military satellites, for example, the high-thrust, lower-Isp mode would be used for LEO-to-GEO tran

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