Abstract:
A pulse plasma thruster ( 50 ) utilizes a vapor producing solid ( 54 ) and a micro-sized heater ( 52 ) to produce a high pressure vapor that is directed into an ignition chamber ( 58 ) and to a thrust discharge chamber ( 70 ). The thrust discharge chamber ( 70 ) comprises two oppositely disposed electrode plates ( 72, 74 ) and oppositely disposed fuel propellants sources ( 60, 62 ). The passageway ( 56 ) leading from vapor producing solid ( 54 ) to the thrust discharge chamber ( 70 ) is configured to permit uniform feeding of the vapors to the thrust discharge chamber ( 70 ). A pair of electrode terminals ( 82, 84 ) extend from the electrode plates ( 72, 74 ) and through a housing ( 88 ). A power source ( 100 ) is coupled to the terminals ( 82, 84 ) and provides the ignition signals necessary to cause a spark and a breakdown to a useful plasma arc by controlling the voltage-current shape of the ignition signal.

Description:
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 2  H 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. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     Other advantages of the invention including specific embodiments are understood by reference to the following detailed description taken in conjunction with the appended drawings in which: 
     FIG. 1 illustrates a prior art liquid thruster that uses a high voltage power supply to create thrust; 
     FIGS. 2 a  and  2   b  illustrate embodiments of a Pulsed Plasma Thruster (PPT) according to the invention; 
     FIG. 3 is a cross section view of the ignition chamber of the PPT of the invention; 
     FIGS. 4 a  and  4   b  illustrate the use and operation of the variable power processing unit that powers the pulsed plasma thruster according to one embodiment; and 
     FIG. 5 illustrate the micro-positioning of the electrode plates of the pulsed plasma thruster according to one embodiment. 
     References in the detailed description correspond to like references in the figures unless otherwise indicated. 
    
    
     DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS 
     The present invention provides a pulsed plasma thruster (PPT) that can be used as a rocket engine for small spacecraft. The PPT operates on a pulse basis where a spark is created at low voltage via the use of small separations of electrodes using micro-electromechanical systems (MEMS) technology, independent introduction of vapor from solids, and electrodes which are slightly radioactive and specially shaped. The spark is transferred to an arc via use of a power supply with three output sections. The arc creates a plasma consisting of constituents of PTFE, which is ablated by the arc, and the vapors from the solids. 
     Referring to FIG. 2 a , a pulsed plasma thruster (PPT) according to one embodiment of the invention is shown and denoted generally as  50 . The PPT  50  includes a heater  52  or other means of generating heat that is small enough to accommodate the framework of a small spacecraft. In one embodiment, the heater  52  is a micro-sized heater based on micro-electromechanical systems or MEMS technology. 
     The heater  52  is placed adjacent a subliming solid  54 . The purpose of the subliming solid  54  is to provide a vapor source so that, in combination with the heater  52 , the solid  54  generates a gas flow that assists ignition of an initial plasma arc in the spark region of the thruster  50 . Thus, the heater  52  increases the temperature of the subliming solid  54  which, in turn, generates vapor. The vapor flows through a screen  56  and into an ignition section  58  of the thrust discharge chamber  70  where a spark partially ignites the solid fuel propellant  60  as well as some of the subliming solid  54  that has been vaporized. The action of the subliming solid  54  and resulting vapor, coupled with the screen  56  and configuration of the ignition section  58  assist in igniting a spark that creates a useful plasma arc. 
     In general, the subliming solid  54  has the characteristic of being able to produce a vapor when heated. A low sublimation temperature of the solid  54  is desired so a large quantity of vapor gas is generated for relatively small incremental changes in temperature. This reduces the heat generating requirements of the heater  52 . While some gases provide better ignition sources than others, the requirement that the solid  54  produce easily ionized vapor restricts selection of the material to certain compounds. Candidates include carbonates (X(HCO 3 ) and carbamates (S(CO 2 NH 2 )) which sublime into NH 3 , CO 2 , and H 2 O. The use of subliming solids enables independent addition of vapor into the PPT, eliminates the requirements for valves and seals, and assures long term compatibility with space environments. 
     FIG. 3 is a cross section of the PPT  50  taken along line  3 — 3  of FIG. 2 a  and illustrating the arrangement of the ignition section of the thrust discharge chamber  70  in greater detail. As shown, the screen  56  contains a plurality of holes  80  which are spaced and sized to provide optimum feeding of vapor from the subliming solid  54  into the thrust discharge chamber  70 . The number of holes  80  depends on the size of the ignition section  58  and the requirement that plasma in the ignition chamber must be allowed to enter the chamber that holds the solid  54  Thus, the sizing, diameter and quantity of the holes  80  is influenced by the specific configuration of the PPT  50 . 
     Preferably, the velocity of the vapor into the ignition section  58  is kept relatively low. In general, many small holes are more effective than a few big holes. Also, the screen  56  is designed to separate the solid from the thrust discharge chamber  70  so that sparks and/or plasma does not interact with solid  54 . 
     As shown, the thrust discharge chamber  70  is comprised of the two oppositely disposed electrode plates  72  and  74  and two fuel propellants  60  and  62 . The fuel propellants  60  and  62  are preferably PTFE based, although other fuel sources may be utilized. In one embodiment, MEMS based micro-heaters (not shown) are embedded in the fuel propellents  60  and  62  and their temperature varied to control the amount of PTFE ablated and to provide more control of the impulse generated by the PPT  50 . In another embodiment, solids  54  are placed along the thrust discharge chamber  70  and nozzle  90 . The solids  54  contain micro-heaters which are independently controlled to allow the introduction of vapor into the thrust discharge chamber at optimum locations and times during the firing cycle. This vapor provides additional control of the efficiency and the impulse bits generated by the PPT  50 . 
     Referring again to FIG. 2 a , the PPT  50  also includes a set of electrode plates  72  and  74 . The electrode plates  72  and  74  correspond to the anode and cathodes of the PPT  50 , respectively. As shown, the distance “d” corresponds to the spacing between the electrode plates  72  and  74 . In one embodiment, the distance “d” between the electrode plates  72  and  74  is 50 micrometers or less. Additionally, the electrode plates  72  and  74  are positioned so that they are evenly displaced about the central axis “x” running through the thrust discharge chamber  70  of the PPT  50 . 
     An advantage of the PPT  50  is the ability to create a reliable breakdown within the thrust discharge chamber  70  using low levels of power. This is achieved, in part, by keeping the spacing “d” between the electrode plates  72  and  74  small so that a spark is more efficiently generated and ignition is achieved using less spark energy. Recent advances in MEMS technology enables the manufacture of small clearances between the electrode plates  72  and  74 . Thus, the fact that the PPT  50  incorporates MEMS technology provides a PPT  50  suitable for space missions where power is limited. 
     According to various embodiments, the electrode plates  72  and  74  are spaced anywhere from 1 micrometer to 50 micrometers apart. In general, the closer the electrode plates  72  and  74  are spaced, the lower voltage is required to a ignite a breakdown. 
     Coupled to the electrode plates  72  and  74  are corresponding electrode terminals  82  and  84  that extend through an insulating layer  86  and the housing  88 . The electrode terminals  82  and  84  are used to deliver the ignition voltage to the thrust discharge chamber  70 . The insulating layer  86  extends substantially over the thrust discharge chamber  70  and the thrust nozzle  90 . As is known to those of ordinary skill in the art, the insulating layer  86  can be configured to increase the local field strengths existing between electrode plates  72  and  74 . 
     A disadvantage of prior art thrusters is that they require very high ignition voltages to operate. For example, the thruster  10  requires a DC supply anywhere from 2000 volts to 8000 volts. Such high voltages have been used in PPTs for a long time since they result in greater thrust. The present invention contemplates the use of voltages less than 300 volts. In one embodiment, the spacing of the electrode plates  72  and  74  is such that 50 volts is sufficient to create suitable thrust levels. This permits the PPT  50  to be utilized in typical satellite applications where 50 volts is commonly available. 
     FIG. 2 b  illustrates another configuration of the PPT  50  according to the invention. Specifically, the PPT  50  is shown equipped with a means of adjusting the angle of the thrust nozzle  90  with respect to central axis “x”. The hinges  92  and  94  are provided for this purpose although other means of achieving the same function can be employed. In this way, the PPT  50  becomes a fuel dynamic device since the angle of the thrust nozzle  90  has some effect on the amount of fuel utilized for certain levels of thrust. 
     Referring to FIG. 4 a , therein is shown the PPT  50  driven by a power source  100  with terminals  102  and  104  coupled to electrode terminals  82  and  84 , respectively. In general, the power source  100  is capable of producing multiple volt-ampere signal forms that effect the shape and magnitude of the ignition signal used to spark the vapors in the ignition section  58 . 
     In one embodiment, the power source  100  comprises a flexible power processing unit that operates in the three segments: an open circuit to constant voltage segment, a constant voltage segment, and a constant current segment. The three segments are illustrated in the graph of FIG. 4 b.    
     The open circuit voltage, Vo, from the power source  100  is applied to the electrodes. The vapor from solid  54  is also introduced into the ignition section  58  of the thrust discharge chamber. A spark occurs in the ignition section  58 . During the next segment, the voltage decreases to the constant voltage section of the power source  100  at current Ic. The current then increases at a constant voltage, Vc, to a constant current section where the current is held constant at Io. The values of Vo, Vc, Ic, and Io are preset to desired values dependent on the specific design and operating condition of the PPT. Designs of power supplies capable of such outputs are known to those of ordinary skill in the art. 
     With reference to FIG. 5, the PPT  50  is equipped with micro-positioning devices  110  and  112  operably coupled to the electrode terminals  82  and  84 , respectively. The purpose of the micro-positioning devices  110  and  112  is to adjust the positioning and spacing of the electrode plates  72  and  74  with respect to the central axis “x”. Preferably, the micro-positioning devices  110  and  112  are MEMS based so that they fit the framework of a small spacecraft and require only small amounts of power to operate. In this way, the spacing between each electrode plates  72  and  74  can be varied as a function of axial distance from the upstream end of the thrust discharge chamber  70 . 
     While the invention has been described in conjunction with preferred embodiments, it should be understood that modifications will become apparent to those of ordinary skill in the art and that such modifications are therein to be included within the scope of the invention and the following claims.