Patent Number: 062164457
Section: summary

TECHNICAL FIELD The invention relates generally to plasma thrusters and more particularly to a miniature pulsed plasma thruster capable of efficiently generating very small impulse bits at low levels of power and DC ignition voltages. BACKGROUND OF THE INVENTION Space vessels such as spaceships and satellites utilize thrusters to achieve motion in space. A thruster operates on the principle that a force generated in one direction generates an equal force in the opposite direction. By emitting a reaction-mass, a thruster accelerates a spacecraft in the opposite direction. A thruster may be used as a small rocket engine for orbit correction or as the main propulsion of the spacecraft. Older conventional thrusters used chemical propulsion, which utilized liquid and/or solid propellants. Electric thrusters, which accelerate gases by electrical heating and/or by electric and magnetic field forces, can outperform chemical propulsion systems, in part, because of their high specific impulse (Isp) values. Advantages of electric thrusters include high efficiency and performance, low weight, increased spacecraft orbiting lifetimes, reduced overall costs, and a savings in fuel mass. Advances in onboard electric power sources and smaller more efficient electronic devices have expanded the use of electric thrusters in spacecraft applications. Electric thrusters that convert electrical energy into kinetic energy may be grouped into three categories: electro thermal propulsion, electrostatic or ion propulsion, and electromagnetic propulsion. Within the electromagnetic propulsion category is the Pulsed Plasma Thruster (PPT), which accelerates the propellant plasma via interaction with an electric arc. Multiple government and civil entities are developing small and micro sized spacecraft that can benefit from PPTs for space missions. Such spacecraft will require major reductions in thrust levels and/or impulse bits to ensure proper and precise control of the spacecraft. Many missions, in particular those that require significant mission propulsion energies and/or acceleration, will require specific impulses beyond those available from chemical rockets. Because present electric rockets cannot efficiently operate a very low level of power and impulse bits they are not well suited for such missions. While PPTs are at a high state of development, they generally require high levels of voltage and power to initiate the plasma breakdown and are also very inefficient at low powers when operated at values of expelled propellant velocities of interest to space missions. For example, experimental PPTs have been operated at energy levels down to about 2 joules (J) per pulse requiring the use of high voltage charging supplies which can range from 2,000 to 8,000 volts depending on the design. Also, efficiencies of PPTs decrease with decreasing power and presently, are less than 10 percent efficient when operated at values of propellant velocities of interest to space systems. The inefficiencies result in significant increases in power to achieve desired levels of impulse bits. An example of such a thruster is shown in FIG. 1 and denoted generally as 10. The thruster 10 fits into the class of propellant devices that operates using an all gas propellent although an all solid solution could also be utilized. In particular, the thruster 10 utilizes a low atomic weight liquid propellant such as water or monopropellant hydrazine (N.sub.2 H.sub.4) or a mixture of two liquids such as water and hydrazine which is stored in the tank 12 and flows through a conduit 14 leading to an opening 16 that forms the feeding mechanism of the thruster 10. The liquid propellent within the tank 12 may be pressurized by high pressure helium in the tank 20, in a manner well known to those of ordinary skill in the art. The liquid propellent flows through the conduit 14 via the opening 16 and reaches a passage 18 within the thruster 10. The passage 18 leads to a small opening 22 which is sized to provide the correct flow velocity for the liquid propellent and reduce back flow into the passage 18. In the passage 18, the liquid propellent is partially or fully atomized and partially evaporated, so that there is a two phase flow of liquid and gas into the thruster 10. The liquid propellent is disassociated into low atomic weight elemental constituents thereof by an electric discharge that forms a plasma arc within the thruster 10. The liquid gas and plasma flow from an open end 24 of the passage 18 into the thrust nozzle 30 which, as shown, is shaped as a cone or bell having a curved confining surface, to provide high efficiency and conversion of the high pressure plasma into a directed supersonic flow having high momentum. This discharge of plasma is established primarily by the use of a high voltage DC (HVDC) power supply 32 which is coupled to electrodes 34 and 36 of the thruster 10. In particular, the thruster 10 operates when liquid from the tank 12 flows into the passage 18 and a high voltage ignition signal supplied by the HVDC power supply 32 is applied at terminals 34 and 36 at a predetermined frequency, such as 200 pulses per second, for example. This ignition voltage can vary but according to one design ranges from 2,000 volts to 8,000 volts. The ignition signal supplied by the HVDC power supply 32 causes a discharge to be established in the passage 18 between the electrodes 34 and 36 at a time when partially atomized fluid is entering the thrust nozzle 30 through the opening 24. The velocity and mass flow rate of liquid flowing through the passage 18 and the repetition rate and energy of the plasma discharge between the electrodes 34 and 36 are matched to achieve optimum operation. Typically, the HVDC power supply 32 raises the voltage of the thruster 10 until an electrical breakdown occurs between the electrodes 34 and 36. The requirement, however, that the HVDC supply 32 generate high levels of ignition voltages makes the thruster 10 unsuitable for many propulsion applications where small spacecraft are involved. The HVDC supply 32 can be large and not well suited for such applications. Moreover due to its size, the HVDC supply 32 makes it difficult to achieve small and precise maneuvers for some spacecraft missions. For many space mission applications, where small space systems are involved and which require extremely precise control, the use of high power and/or high voltage ignition circuits is impractical. Examples of such missions are those which require extremely precise ephemeris control and those which are otherwise penalized by high thrust, such as missions which require multiple acceleration and deceleration maneuvers. Thus a PPT that is able to efficiently operate without a high voltage ignitor system and at power levels several orders of magnitude less than prior art designs would be advantageous. SUMMARY OF THE INVENTION The present invention is a pulsed plasma thruster (PPT) capable of operating at low levels of power and impulse bits that is suitable for use in space applications where the space system is small and precise control of the spacecraft is required. The PPT of the present invention is capable of delivering reliable ignition of a spark breakdown at DC voltages less than 300 volts with reliable transfer of a spark to a useful plasma arc. The ablation, combustion and acceleration of the Polytetra Fluorethylene (PTFE) fuel propellent is precisely controlled with the use of miniaturized PPT and power processor components. The efficiency of the thruster is increased by the independent introduction of vapor (such a from a subliming solid) at optimal locations and times during the operational cycle. According to one embodiment, disclosed is a PPT having optimally located solids capable of producing high vapor pressures for purposes of enhancing both ignition and efficiency. Heat generating elements, such as micro-heaters, are placed adjacent to the solids and configured to generate heat that causes the solids to sublime. The PPT includes an igniter section that forms a passageway from the solid to a thrust discharge chamber. In one embodiment, the ignition chamber includes a plurality of holes which are sized and spaced for optimally guiding vapors to the thrust discharge chamber for purposes of enabling arc ignitions at low voltages. In one embodiment, solids are also located within the thrust discharge chamber and, via the use of heat generating elements, independently introduce vapors into the thrust discharge chamber in order to enhance PPT efficiency at desired values of propellant velocities. The thrust discharge chamber includes a set of properly spaced and shaped electrode plates which provide for transfer of an initial spark to a useful plasma arc in the gap defined by the electrodes plates. A solid propellent, such as PTFE, is provided within the thrust discharge chamber and arranged so that the plasma arc traveling through the thrust discharge chamber will ablate the PTFE and accelerate the plasma formed from ablated PTFE and the independently introduced vapor from high vapor pressure solids, as used. A power processing unit provides the DC ignition voltage necessary to cause a spark to occur in the gap between the electrode plates. In one embodiment, the power processor unit has a variable output that operates in three segments: an open circuit to constant voltage segment, a constant voltage segment, and a constant current segment. A high vapor pressure between the electrode plates is created when the heat generating means heats the solid to assist in ignition and transition of a spark to a useful plasma arc. Micro-heaters can also be embedded in, or at the edges of, the PTFE propellent and its temperature varied to control the amount of PTFE ablated to provide more control of the impulse generated by the PPT. Micro-heaters embedded in the solids, located in the ignitions section and/or the thrust discharge chamber, independently provide a source of vapor to the thrust discharge chamber to provide additional and independent control of the efficiency and impulse of the PPT. In one embodiment the electrode plates are equally spaced about a central axis through the thrust discharge chamber. In another embodiment, the PPT includes a means of varying the spacing between the electrode plates as a function of axial distance. In an other embodiment, slightly radioactive electrodes are used. In these ways ignition voltages and required power levels are achieved that are several orders of magnitude smaller than those previously obtainable. Also disclosed is a method of operating a pulsed plasma thruster comprising the steps of heating a subliming solid to create a high pressure vapor and directing that high pressure vapor in the direction of a thrust discharge chamber through an ignition chamber. Next, a DC ignition signal is applied to electrodes coupled to the thrust discharge chamber that sparks a breakdown of a fuel propellent and causes a transition of the spark to a useful plasma arc. The DC ignition signal is applied in a way that its shape and magnitude are controlled. In one embodiment the DC ignition signal is controlled in three segments corresponding to an open circuit to constant voltage segment, a constant voltage segment and a constant current segment. The high pressure vapor is directed to the thrust discharge chamber so that pressure is created between two electrode plates. The vapor can be fed uniformly to control ignition and breakdown of the fuel propellent. The spacing between the electrode plates may be adjusted to control the amount of the fuel propellent ablated. A source of ultraviolet radiation may be focused on the vapor to provide additional excitation energy that helps ignite the vapor from the subliming solid. A technical advantage of the invention is the enablement of reliable ignitions at voltages more than an order of magnitude less that previously obtainable. This enables small and light-weight PPTs and power supplies and, therefore, much lighter PPT systems than previously obtainable. Another advantage is the efficient enablement of impulse bits several orders of magnitude less than previously obtainable. This enables the deployment of PPTs suitable for space propulsion applications involving small spacecraft systems and for missions which require extremely precise control of the spacecraft.