Patent Number: 062164457
Section: description

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. 2a, 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.sub.3) and carbamates (S(CO.sub.2 NH.sub.2)) which sublime into NH.sub.3, CO.sub.2, and H.sub.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. 2a 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. 2a, 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. 2b 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. 4a, 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. 4b. 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.