Electric heating – Metal heating – By arc
Reexamination Certificate
2000-03-02
2002-04-16
Walberg, Teresa (Department: 3742)
Electric heating
Metal heating
By arc
C219S121480
Reexamination Certificate
active
06373023
ABSTRACT:
BACKGROUND OF THE INVENTION
(1) Field of the Invention
This invention relates to thrusters, and more particularly to arc initiators for pulsed plasma thrusters for spacecraft.
(2) Description of the Related Art
A background in pulsed plasma thruster (PPT) technology may be found in Cassady, R. Joseph, “Pulsed Plasma Mission Endurance Test”, Air Force Report #AFAL-TR-88-105, August, 1989, the disclosure of which is incorporated herein by reference in its entirety as if set forth at length.
The surge in the use of small spacecraft, especially in deep space constellations, such as ST-3 or Terrestrial Planet Finder (TPF), and in Earth sensing missions, such as EO-1, demands new onboard propulsion solutions. These missions often require a challenging combination of fine impulse control, high specific impulse and maximum thrust for minimum power. PPT's bring proven flight heritage, inert storage, very small impulse bits and high specific impulse for small, low power spacecraft. PPT's also present the option for providing an all-thruster attitude control system (ACS) for any size spacecraft, eliminating the need for wheels and momentum dumping thrusters and resulting in a significant net ACS mass savings.
Typical PPT's are inherently simple, inert and self-contained devices that use an inert solid propellant, typically polytetrafluoroethylene (PTFE), that is ablated and electromagnetically accelerated by an electric arc between two electrodes, very similarly to a plasma “rail gun”. An anode is spaced apart from the cathode (e.g., by an exemplary distance on the order of an inch in a parallel plate thruster configuration). A power source charges an energy storage device (e.g., a capacitor) to anywhere from one to one thousand joules, although 20 joules is a typical value. This charge places the anode at a potential of about 500-3000 volts above the cathode. A separate spark plug is used to initiate the arc discharge. Once the propellant is ablated and ionized by the arc, it is accelerated between the electrodes under the action of a Lorentz body force.
Several first-generation PPT's have been flown in existing spacecraft. A recent PPT system has a total mass, including thruster, electronics, propellant and propellant feed system of around 5 kg. That system can potentially deliver 15,000 N-s, in impulse bits of a fraction of a mN-s for an input power under 100 W. Input power is usually delivered at 28 V, also enhancing the integrability with most spacecraft busses. A PPT system with 8 thrusters an order of magnitude smaller is presently being developed in conjunction with Primex Aerospace Company for the University of Washington Dawgstar satellite, a 10 kg-class spacecraft.
Despite the very promising flight history of PPT's and recent dramatic improvements in PPT design, there are key aspects of the PPT for which improvement would lead to significant reductions in mass, complexity and integration costs. One such area that could hold the key to considerably more widespread usage of PPT's is in its discharge initiation.
Existing methods for initiating (igniting) a PPT discharge present cost and reliability concerns. A common configuration places an annular semiconductor spark plug in the thruster cathode. A spark plug design consisting of a set of coaxial electrodes separated by a ceramic bushing, one end of which is fused with semiconducting material, has been used successfully for many years to ignite PPTs in conjunction with circuitry designed to cause this plug to form a spark under vacuum conditions. An energy storage device (e.g., a capacitor), separate from the main energy storage capacitor, is charged to on the order of half a joule. When coupled by a high voltage switch to the spark plug, this smaller energy storage capacitor induces a flashover between the electrodes of the plug. A basic discussion of flashover and theorized flashover mechanisms is discussed in H. Craig Miller, “Surface Flashover of Insulators”, IEEE Transactions on Electrical Insulation, Vol. 24 No. 5, October 1989, Pages 765-786, the disclosure of which is incorporated herein by reference in its entirety as if set forth at length. See also, Palumbo, D. J., “Solid Propellant Pulsed Plasma Propulsion System Development for N-S Stationkeeping”, AIAA Paper 79-2097, 14th IEPC, Princeton, N.J., 1979.
The spark across the spark plug produces electrons which are drawn toward the thruster anode. As the electrons are drawn to the anode, they come into contact with propellant (such as along the exposed surface of a fuel bar) causing ionization of and electron release from the propellant and initiating the main arc between the thruster anode and cathode. The energy released in the main arc may be approximately one hundred times greater than the energy released in the arc across the spark plug.
Existing spark plugs as well as some of the associated high voltage equipment (e.g., insulated gate bipolar transistors (IGBT)) present particular reliability risks. In addition to unexpected failure, existing spark plugs have inherent lifetime limitations. The plugs can easily be the life limiting component for the entire PPT system, providing less than one million pulses under some circumstances, up to a maximum proven life of ten million pulses for a known configuration. Future uses of PPT's will require twenty-forty million pulse lifetimes or greater. Aside from total failure of the spark plugs, performance decay over the functional lifetime of the spark plug can produce associated changes in thruster performance. By way of example, a new spark plug may have a breakdown voltage of as low as about 200 volts. Over its lifetime, the breakdown voltage will increase, for example to about 2,000 volts. Another performance concern is the more random shot-to-shot variability of PPT thrust pulses. Studies have shown that much of this variability can be correlated with variability in the location of the discharge initiation spark, which, due to the annular design of existing spark plugs, is relatively wide. For smaller PPT designs the impact of this problem becomes more significant.
Another problem associated with PPT's is electromagnetic interference (EMI). Studies have shown that a significant fraction of the EMI signature of a PPT is due to the spark event, which is a comparatively high frequency phenomenon relative to the main arc discharge (further into the frequency range of concern for EMI).
Another issue is weight. The circuitry utilized to generate the fast, high voltage, spark of the spark plug can occupy approximately one-half of the electronics board area for a PPT. By way of example, an exemplary circuit includes an 800 volt source and a 3:1 step up to achieve the necessary spark plug breakdown voltages anticipated over the plug's lifetime.
BRIEF SUMMARY OF THE INVENTION
The invention seeks to initiate arc discharge by preferably introducing electrons very close to the propellant. This may be achieved by thermionic emission of electrons. The thermionic emission can be provided via relatively low voltage circuitry which can reduce weight and EMI as well as cost and, potentially, power consumption. Thruster life may be significantly improved via use of components which are not subject to significant erosion and/or use of components which, although subject to erosion, are replenished such as in the self-feeding mounting of a propellant bar.
Accordingly in one aspect the invention is directed to a pulsed plasma thruster comprising a pair of electrodes being an anode and a cathode spaced apart from the anode. A voltage source applies a voltage between the cathode and the anode to positively charge the anode relative to the cathode; a solid propellant bar extends longitudinally and is held for progressive advancement in a downstream longitudinal direction to a gap between the cathode and anode. An initiator initiates arc discharge between the anode and cathode by inducing thermionic emission of electrons, which electrons are drawn toward the anode and tend to induce ionization of ma
Cassady Robert J.
Hoskins William A.
General Dynamics (OTS) Aerospace, Inc.
Slate William B.
Van Quang
Walberg Teresa
Wiggin & Dana
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