Abstract:
A spacecraft attitude and altitude control system utilizes sets of three pulsed plasma thrusters connected to a single controller. The single controller controls the operation of each thruster in the set. The control of a set of three thrusters in the set makes it possible to provide a component of thrust along any one of three desired axes. This configuration reduces the total weight of a spacecraft since only one controller and its associated electronics is required for each set of thrusters rather than a controller for each thruster. The thrusters are positioned about the spacecraft such that the effect of the thrusters is balanced.

Description:
This is a continuation application of provisional application U.S. Ser. No. 60/076,435, filed Mar. 2, 1998 which is incorporated in its entirety herein. 
    
    
     The U.S. Government has a paid-up license in this invention and the right, in limited circumstances, to require the patent owner to license others on reasonable terms as provided for by the terms of contract number NAS3-27570 awarded by NASA. 
    
    
     BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     This invention relates to pulsed plasma thrusters for spontaneously adjusting the attitude or altitude of a spacecraft. More particularly, this invention relates to a single controller that controls the magnitude and direction of thrust in a thruster set. 
     2. Description of the Art 
     Present spacecraft attitude and altitude control systems use a combination of rotating wheels with either thrusters or magnetic torque rods to compensate for changes in wheel speed when the spacecraft attitude is adjusted. The wheels are used to absorb angular momentum or to generate rotation of the spacecraft in response to disturbance torque imparted to the spacecraft. Wheels have been used because a system utilizing only thrusters was considered to be limited by the amount of propellant that may be carried on board. 
     Thrusters have not been used to their fullest potential because even a very small control thrust generated by typical thrusters is too large and such a thrust caused the spacecraft to be overwhelmed. This required compensation measures to correct the motion of the spacecraft. This problem in spacecraft control is due to the fact that many designs provide thrust in two directions or alternatively, cant the electrode to provide thrust at an angle. These designs are not desirable since thrust is provided in two components and thus, requires firing a second thruster to compensate for any undesired torque or translations produced as a result of the components of the first thrust. 
     U.S. Pat. No. 4,848,706 (Garg et al.) discloses a spacecraft attitude control system using coupled thrusters. The thrusters provide cross coupling torque so that as one thruster is fired; a second thruster is fired thereby providing a balancing force. 
     U.S. Pat. No. 5,207,760 (Dailey et al.) discloses a multi-megawatt pulsed inductive thruster. A gas is discharged against an inductor. Each thruster has an associated capacitor to discharge by a trigger generator after a puff of propellant reaches an inductor. 
     U.S. Pat. No. 5,339,623 (Smith) discloses singly fueled multiple thrusters simultaneously energized by a common power supply. This technique allows a single power supply for multiple thrusters but does not switch from one thruster to a second thruster. 
     U.S. Pat. No. 5,439,191 (Nichols et al.) discloses a railgun thruster, which provides attitude control via three perpendicular axes. This patent discloses that each thruster has an associated power source. 
     Thus, there is a need for an improved thruster that provides a single control unit and power source for a plurality of thrusters such that the proper magnitude of thrust is provided to maintain the desired direction of the spacecraft. This type of thruster would eliminate the need for additional thrusters to be fired. 
     BRIEF DESCRIPTION OF THE INVENTION 
     Accordingly, it is an object of the present invention to provide a spacecraft control system that adjusts the attitude and/or altitude of a spacecraft such that the orbit of the spacecraft is not disturbed. 
     It is also an object of the invention that a single controller controls sets of thrusters of a spacecraft. 
     A third object of the invention is the placement of thruster sets about a spacecraft to provide control in all three orthogonal directions. 
     It is a feature of this invention that an attitude control system includes sets of three axial pulsed plasma thrusters, coupled to a single controller that selectively provides thrust along a desired axis. A low inductance stripline couples the controller to a set of discharge electrodes on each axial thruster. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     FIG. 1 shows a three axis thruster set. 
     FIG. 2 shows an exploded view of one of the axial thrusters in a three axis thruster set. 
     FIG. 3 shows a planar view of the three axis thruster stripline. 
     FIG. 4 shows a stripline design for a three axis pulsed plasma thruster that is configured around a cylindrical capacitor. 
     FIG. 5 shows placement of the three axis thruster modules around a spacecraft. 
     FIG. 6 shows the charging circuitry for a controller. 
     FIG. 7 shows the firing circuitry for the thruster. 
     FIG. 8 shows an embodiment of a three-axis thruster set about a single controller. 
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Thrusters are used to control spacecraft in order to maintain a desired orbit and orientation. The present invention uses a single controller in conjunction with axial thrusters to adjust the attitude and/or altitude of a spacecraft. Three axial thrusters are electronically coupled to a single controller, which determines the direction and magnitude of thruster firing. The three axial thrusters form a thruster set, which may be configured so that there is a thruster in the X-axis, Y-axis and Z-axis. Each axial thruster has discharge electrodes and a propellant bar. The use of a single power source for pulsed plasma thrusters is advantageous because there is a substantial reduction in the cost of manufacture and weight of the device. The inventive configuration eliminates the need to compensate for thrust in an undesired direction because only thrust in a single direction is provided by the single controller and therefore, no undesired thrust components are introduced. 
     FIG. 1 shows the three axis thruster set  10  that is comprised of a first axial thruster  120  along the X-axis, a second axial thruster  130  along the Y-axis and a third axial thruster  140  along the Z-axis. Each axial thruster  120 ,  130  and  140  has a propellant bar associated with it. Propellant bar  180  is associated with the X-axis thruster  120 , propellant bar  181  is associated with the Y-axis thruster  130  and propellant bar  182  is associated with the Z-axis thruster  140 . 
     The controller  136  provides a command to a desired axial thruster thereby providing thrust in a specific direction. The controller  136  provides control over all three axial thrusters that make up the thruster set  10 . 
     The controller  136  includes: a capacitor (shown in FIG. 2 as element  138 ) for energy storage and an on-board power converter, which takes power from the spacecraft in order to charge the capacitor. The capacitor receives power at a low rate, ranging from about 0.01 to 1.0 Watts and more typically approximately a tenth of a Watt per a time interval that ranges between 0.01 and 3 seconds and more typically about 0.25 seconds. The capacitor then discharges the accumulated power in a very short time interval, approximately 1 to 10 microseconds and more typically about 5 microseconds. This rapid discharge provides the necessary high power for accelerating the plasma, thus imparting thrust to the spacecraft. 
     The controller  136  determines the direction of thrust by a propellant fuel bar. These propellant fuel bars,  180 ,  181 ,  182  melt a material such as synthetic resinous fluorine containing polymers, for example polytetrafluoroethylene, causing it to vaporize. The high current discharge then ionizes the vaporized propellant and accelerates it down the electrodes, thereby imparting thrust to the spacecraft. 
     As shown in FIG. 2, a single controller  136  and a capacitor  138 , are used in conjunction with the present system. The controller  136  takes low voltage power from the spacecraft and provides high voltage outputs to charge an energy storage capacitor  138 , and provides a high voltage pulse to fire one of a plurality of sparkplugs. For discussion purposes, only a single axial thruster will be described. Each axial thruster has a single sparkplug associated with it, i.e., sparkplug  193 , with axial thruster  120 . The energy storage capacitor  138  utilizes anode stripline  150  and cathode stripline  160  for connecting the capacitor  138  to the anode discharge electrode  170 ; and the cathode discharge electrode  171  and a propellant fuel bar  180 . The voltage across the capacitor  138  is the voltage across the electrodes  170 ,  171  of the X-axis thruster as well as the voltage across the electrodes of the Y-axis and Z-axis thrusters. This is possible because the anode stripline has very low resistance. 
     The discharge electrodes  170 ,  171  are associated with X-axial thruster  120 . The Y-axial thruster  130  and Z-axial thruster  140  also have associated discharge electrodes (not shown). The anode stripline  150  provides a path for current from the capacitor  138  to the anode discharge electrode  170  when an axial thruster is fired. The anode stripline  150  functions as an extension of the capacitor as well as a high current carrying conductor with a magnitude that ranges between 10,000-40,000 amperes. The cathode stripline  160  is coupled to the capacitor such that a connection between the capacitor  136  and the cathode discharge electrode  171  is provided. The anode and cathode striplines must be insulated to withstand a voltage in excess of 2,000 volts. 
     The sparkplug  193  has associated electronics in the controller  138  to implement the firing of an associated axial thruster. 
     As shown in FIG. 3, the controller  136  employs the anode stripline  150  and the cathode stripline  160  to provide a current path from a power source to the electrodes of a particular axial thruster  120 ,  130  or  140 . The axial thrusters are arranged so as to provide thrust in all three orthogonal directions. This requires both the anode stripline  150  and the cathode stripline  160  to be angled 90 degrees at the locations indicated by the dashed line  3 - 1  in FIG.  3 . This eliminates the need for counter-balancing thrusters. Propellant fuel bars provide the required material to produce thrust. 
     Cathode stripline  160  ranges in length from approximately 8 to 15 inches. 
     FIG. 4 shows a side view of the anode stripline  150  for all three thrusters. This anode stripline, which is comprised of a thin copper sheet having very low inductance, has a length of approximately eight to fifteen inches and provides for current conduction between the energy storage capacitor  138  and the discharge electrodes  170 ,  172 ,  174  for all three axes. A similar configuration is-employed for the cathode stripline  160 . Discharge electrode  170  is the discharge electrode used in conjunction with the X-axial thruster. Discharge electrodes  172  and  174  are used in conjunction with the Y and Z axial thrusters respectively. Discharge electrode  170  has associated thin copper sheets  478  and  476  to provide conduction to discharge electrode  170 . In a similar fashion, discharge electrode  172  has associated conductors  471  and  472 , and discharge electrode  174  has associated conductors  474  and  476 . 
     FIG. 5 shows that the propellant bars  180 ,  181 ,  182  may be configured to be symmetric around the controller  136  to minimize inductance of the striplines and maintain high efficiency when they are fired. Also, as shown in FIG. 5, the thruster sets  10  may be positioned on a spacecraft  12  such that the orbit of the spacecraft is maintained either through a translation or a rotation maneuver when an axial thruster  120 ,  130  or  140  is fired. This is a desired configuration since it provides complete stability for the spacecraft with a minimum number of pulsed-plasma thruster units. 
     Pulsed plasma thrusters overcome the limitation of the amount of propellant that may be carried on board a spacecraft because they achieve high specific impulse through the use of electrical energy. Unlike other electric thrusters, pulsed plasma thrusters are pulsed devices that use an energy storage capacitor that allows operation at a much lower average power level. Pulsed plasma thrusters such as thruster set  10  take in power at a low rate and discharge it very rapidly thereby providing high instantaneous power to produce thrust at a high specific impulse. The specific impulse (Isp) is desired to have a large magnitude for sustained propulsive maneuvering. The specific impulse for these axial thrusts is approximately 1000 sec., which is a measure of the thrust per unit of propellant used. Pulsed plasma thrusters provide very precise impulses that will not overwhelm the spacecraft. The minimum impulse bit (Ibit) typically achievable ranges in magnitude from 50 to 100 μNs. This combination of a minimum Ibit and high specific impulse bit provides an efficient thruster for spacecraft maneuvers. 
     FIG. 6 shows charge conversion circuitry  650  for the controller. Circuit  650  is representative of the circuitry to fire thrusters in the X and Y directional components. Circuitry to provide thrust in the Z component would include similar features. The purpose of circuit  650  is to charge the energy storage capacitor. This is accomplished by receiving a command input at terminal  610  and a low voltage power input from the spacecraft (typically between 20 and 30 volts) at terminal  620 . The charge conversion circuit  650  outputs a high voltage across the terminals of the capacitor through terminal  630 , thus charging it with stored energy. The conversion circuit  650  also outputs a discharge initiation voltage to initiate firing of a spark plug through output  640 . Output  640  is used to fire the spark plug associated with the X-axis thruster and the spark plug associated with the Y-axis thruster. 
     FIG. 7 shows details of discharge initiation circuitry  640  to cause spark plugs in the X-axis thruster and Y-axis thruster to fire. The control of the firing of a selected axial thruster (x, y, or z) is accomplished by selecting which sparkplug to fire. This is accomplished via a command signal from the spacecraft. The command to fire the X-axis spark plug is received by input  720  of the discharge initiation circuit  640  and outputs a command to X-axis spark plug at terminal  730 . The command to fire the Y-axis spark plug is received by input  740  and outputs a command to the Y-axis spark plug at terminal  750 . 
     FIG. 8 shows the three axis thruster set  10  with each axial thruster  120 ,  130  and  140  placed 120° from each other. This configuration shows that the-axial thrusters  120 ,  130 ,  140  may be configured to enable thrust in each of three orthogonal axes of the spacecraft when fired. 
     While the preferred embodiments of the present invention have been illustrated in detail, it should be apparent that modifications and adaptations to those embodiments may occur to one skilled in the art without departing from the scope of the present invention as set forth in the following claims.