Power plants – Reaction motor – Electric – nuclear – or radiated energy fluid heating means
Reexamination Certificate
2002-07-09
2004-08-03
Yu, Justine R. (Department: 3746)
Power plants
Reaction motor
Electric, nuclear, or radiated energy fluid heating means
C060S253000, C060S256000, C060S247000
Reexamination Certificate
active
06769241
ABSTRACT:
FIELD OF INVENTION
The invention relates to propellant designs in Micro Pulsed Plasma Thrusters, and particularly to techniques for bundling propellant modules and for using a two-stage discharge process to increase MicroPPT propellant throughput, and decrease the output voltage required from the power-processing unit.
BACKGROUND OF INVENTION
Micro Pulsed Plasma Thrusters (MicroPPTs) are small thrusters used for propulsion and attitude control on spacecraft. The MicroPPTs are designed to provide spacecraft propulsion at low power levels (<20 W), in small reproducible impulse bits of thrust, and be contained in a very small lightweight package (<200 grams). For 100-kg class microsatellites, MicroPPTs can provide propulsive attitude control and at least partial stationkeeping propulsion. For 25-kg class or smaller satellites, MicroPPTs can provide all primary propulsion, as well as stationkeeping and attitude control.
While MicroPPTs retain some design similarity with standard PPTs, they are fundamentally different in critical areas that enable the required reductions in mass, size, and power. For example, the LES-8/9 PPT, designed and flight tested in the 1970s and shown schematically in FIG.
1
and in the photograph of
FIG. 2
, a DC—DC converter charges an integrated capacitor from a 28-V spacecraft bus to 1500 V. Then, a second DC—DC converter supplies 600 V to a smaller capacitor in the trigger circuit.
Modern flight units operate in a similar regime. The PPT discharge is initiated by a TTL pulse applied to a semiconductor switch in a trigger circuit. The trigger discharge fires a sparkplug embedded in the cathode, providing enough surface ionization or seed plasma to initiate a main discharge across a Teflon™ propellant face. The solid propellant is converted to vapor and partially ionized by the electric discharge. Acceleration is accomplished by a combination of thermal and electromagnetic forces to create usable thrust. As the propellant is consumed, over some 17 million discharges, a negator spring passively feeds the 25-cm-long propellant bar forward between the electrodes.
Although such systems have been well-tested, the required integration of two separate discharge circuits and two separate DC—DC converters dramatically increases the size of the discharge capacitor and the complexity of the PPT propulsion system. MicroPPTs reduce the need for these duplicative and complicated triggering electronics schemes.
Accordingly, the primary difference between the standard PPT and the MicroPPT lies in the electronics. By reducing the size of the thruster, it becomes possible to initiate the discharge by simple over-voltage breakdown at the propellant face, rather than requiring a separate spark trigger. This eliminates one of the two separate discharge circuits and DC—DC converters required in a conventional PPT, and dramatically reduces the size of the discharge capacitor.
However, although the electronics for these MicroPPTs are much improved, the MicroPPT propellant throughput is fundamentally limited. Adding propellant requires the use of a longer propellant rod or a larger diameter propellant rod. A longer rod becomes impractical for spacecraft integration after approximately 12 inches from a geometric packing perspective. Furthermore, for designs where the propellant recesses back into the electrode shell over time, viscous losses with the wall may decrease thrust. In addition, larger diameter propellants are problematic for the DC—DC converter since higher voltages are needed to initiate the surface discharge.
For example, Laboratory tests have defined two critical design criteria for MicroPPT propellants. First, the voltage required to initiate the surface breakdown increases with propellant diameter since the path length between the inner and outer electrodes increases. Second, the average energy per discharge pulse must increase with propellant face area in order to facilitate complete decomposition of the propellant surface. Insufficient energy leads to charring at the propellant face, which dramatically increases the voltage required to initiate the discharge. This voltage increase is considered a failure mechanism for the MicroPPT.
Accordingly, a need exists for an improved propellant design with large cross-sectional area that also has a lower surface breakdown voltage.
SUMMARY OF INVENTION
The present invention relates to propellant designs in Micro Pulsed Plasma Thrusters, and particularly to techniques for creating propellant modules capable of increasing MicroPPT propellant throughput, and decreasing the output voltage required from the power-processing unit.
In one embodiment, several modules of relatively smaller diameter are fabricated into a bundle. Such devices are referred to as “Bundled” propellant designs. In such a “bundled” embodiment the MicroPPT could be designed to use one propellant rod until consumed, and then passively switch to the next one.
In another embodiment, a single module of relatively large cross-section is utilized and ignited using a two-stage ignition process. Such devices are referred to as “Two-Stage” propellant designs.
In yet another embodiment, the invention is directed to thrusters utilizing such propellant modules. The propellant modules of the current invention may be utilized in either MicroPPT or Pulsed Plasma Thruster (PPT) designs. In addition, the propellant modules of the current invention may be utilized in any of the MicroPPT designs listed above and only differ in how the high voltage pulse is electrically generated and then applied to the propellant.
REFERENCES:
patent: 5847315 (1998-12-01), Katzakian, Jr. et al.
patent: 5924278 (1999-07-01), Burton et al.
patent: 6153976 (2000-11-01), Spanjers
patent: 6216445 (2001-04-01), Byers et al.
patent: 6269629 (2001-08-01), Spanjers
patent: 6373023 (2002-04-01), Hoskins et al.
Altman et al., “Chapter 7: Hybrid Rocket Propulsion Systems,” Published in Space Technology Series: Space Propulsion Analysis and Design, McGraw-Hill, 1995, pp. 365-441.
Ashby et al., “Quasi-Steady-State Pulsed Plasma Thrusters,” AIAA Journal, vol. 4, No. 5, May 1966, pp. 831-835.
Aston et al., “Ignitor Plug Operation in a Pulsed Plasma Thruster,” Journal of Spacecraft and Rockets, vol. 19, No. 3, May-Jun. 1982, pp. 250-256.
Barber et al., “Microthrusters Employing Catalytically Reacted N2-O2-H2 Gas Mixtures, Tridyne,” Journal of Spacecraft and Rockets, vol. 8, No. 2, Feb. 1971, pp. 111-116.
Bartoli et al., “A Liquid Caesium Field Ion Source for Space Propulsion,” J. Phys. D: Applied Phys., vol. 17, No. 12, Dec. 1984, pp. 2473-2483.
Bassner et al., “The Design of RITA Electric Propulsion System for Sat 2 (Artemis),” AIAA/DGLR/JSASS 21st International Electric Propulsion Conference, Jul. 1990, Orlando, Florida, AIAA Paper No. 90-2539, pp. 1-7.
Bayt et al., “A Performance Evaluation of MEMS-based Micronozzles,” 33rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Jul. 1997, Seattle Washington, AIAA Paper No. 97-3169, pp. 1-10.
Beattie et al., “Flight Qualification of an 18-mN Xenon Ion Thruster,” AIAA/AHS/ASEE Aerospace Design Conference, Feb. 1993, Irvine, California, AIAA Paper No. 93-1085, pp. 1-7.
Berkopec, “Performance of Two Subliming Solid-Propellant Thruster Systems for Attitude Control of Spacecraft,” NASA Technical Note No. D-3841, Washington, D.C., Feb. 1967, pp. 1-15.
Blandino et al., “Propulsion Requirements and Options for the New Millennium Interferometer (DS-3) Mission,” 34th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Jul. 1998, Cleveland, Ohio, AIAA Paper No. 98-3331, pp. 1-11.
Bromaghim et al., “An Overview of the On-Orbit Results From the Esex Flight Experiment,” 35th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Jun. 1999, Los Angeles, California, AIAA Paper No. 99-2706, pp. 1-13.
Brown, “Spacecraft Propulsion, Chapters 1 and 3,” AIAA Education Series, (1995), pp. 1-8 and 25-54.
Bryant et al., “Planetary Lander Vehicles Utilizing LEAP Technology,” 30th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Jun. 1994, Indianapolis, In
Schilling John
Spanjers Gregory G.
White David
Christie Parker & Hale LLP
Rodriguez William H.
W. E. Research LLC
Yu Justine R.
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