Abstract:
A solar powered spacecraft power system including a solar photovoltaic array, an electric propulsion system connected directly to the solar photovoltaic array in parallel with the spacecraft power system; the electric propulsion system including a Hall effect thruster, a thruster power supply for driving the thruster; a sensor for sensing the power output of the solar array and a controller responsive to the power output of the solar array and configured to periodically adjust an operating parameter of the thruster to operate the thruster at the maximum available output power of the solar array including comparing a previous solar array output power level with a later solar array output power level, and incrementing the operating parameter with a positive value if the later is greater and with a negative value if the later is lesser; a solar powered spacecraft power system including a solar photovoltaic array, an electric propulsion system connected directly to the solar photovoltaic array, and a power management and distribution system connected to the solar photovoltaic array for distributing power to one or more bus loads and payloads.

Description:
RELATED APPLICATIONS 
     This application claims benefit of and priority to U.S. Provisional Application Ser. No. 61/277,766 filed Sep. 29, 2009 under 35 U.S.C. §§119, 120, 363, 365, and 37 C.F.R. §1.55 and §1.78 incorporated herein by this reference. 
    
    
     FIELD OF THE INVENTION 
     This invention relates to an improved solar powered spacecraft power system and to spacecraft electric propulsion for a Hall effect thruster. 
     BACKGROUND OF THE INVENTION 
     In conventional solar powered spacecraft, power is generated by a solar photovoltaic array (SP) and flows into a power management and distribution (PMAD) system. From there it is distributed to all loads: the bus loads, the payload(s) and in electric propulsion (EP) spacecraft, the EP load such as Hall effect thrusters. A PMAD system is a large, heavy, complex collection of circuits including e.g. filters, batteries, DC/DC converters, isolation circuits and voltage regulators. The EP load may use the largest share of the power when it is operating. This means that the PMAD system must have the capacity to process all the power produced by the SP including power for the EP even though significant amount of the time the PMAD system is servicing only the bus load(s) and payload(s). Thus the PMAD system has to be quite large to supply the required power to the EP on demand and also has the have the means to dissipate (as heat) excess power generated by the SP when the EP is making no demand. The extra size and weight required to perform both these functions is critical in spacecraft design and operation. 
     There is yet another shortcoming associated with current PMAD systems. The PMAD system may contain peak power tracking or solar array shunt circuitry in order to optimize the power provided to operate the EP and other loads. At the beginning of life (BOL) the SP provides greater peak power than at the end of life (EOL). The PMAD system peak power tracking unit (operating in continuous mode or discrete steps) lowers the voltage as the SP ages. However, the EP must always remain at a power level that is below the temporary maximum power, never at the maximum power, in order to preserve an operational stability margin to account for the unknown and unpredictable aging of the SP. Assume, for illustration sake that there are no other loads on the SP: the EP operating power point dictates the output power of the array which must be below the peak power point to provide for the operational stability margin. This means that the EP thruster does not get all the power it could get and hence the thrust is reduced and spacecraft transit/maneuver time is extended. In one prior attempt to improve solar powered spacecraft power system the power processor unit conditions the power output from the solar array and regulates the bus voltage and current so as to provide an output current applicable to be used on an arcjet thruster. U.S. Pat. No. 5,604,430. In another approach a power control circuit employs a multiplier, differentiator, detector, phase comparator and integrator to effect a ramp generator to produce minor variations in a beam current reference signal to an ion thruster. U.S. Pat. No. 4,143,314. 
     SUMMARY OF THE INVENTION 
     In accordance with various aspects of the subject invention in at least one embodiment the invention presents an improved solar powered spacecraft power system for Hall effect thruster which reduces power losses, size, weight and costs associated with the PMAD system and improves efficiency by operating the thruster(s) at the peak power point of the solar array power output despite aging degradation of the solar array. 
     The subject invention results from the realization that, in part, an improved solar powered spacecraft power system for a Hall effect thruster in various aspects can be achieved by connecting the electric propulsion system directly to the solar photovoltaic array without the PMAD system and further by periodically adjusting an operating parameter of the thruster to operate the thruster at the maximum available output power of the solar array. 
     The subject invention, however, in other embodiments, need not achieve all these objectives and the claims hereof should not be limited to structures or methods capable of achieving these objectives. 
     This invention features a solar powered spacecraft power system including a solar photovoltaic array, an electric propulsion system connected directly to the solar photovoltaic array, and a power management and distribution system connected to the solar photovoltaic array for distributing power to one or more bus loads and payloads. 
     In a preferred embodiment the solar photovoltaic array may include a number of solar panels. The electric propulsion system may include a Hall effect thruster. The electric propulsion system may include an electric propulsion thruster and a thruster power supply interconnected between the electric propulsion thruster and the solar photovoltaic array. The thruster power supply may include a d.c. to a.c. switching circuit connected to the solar photovoltaic array, a first rectifier connected to the electric propulsion thruster and a transformer with its primary interconnected with the switching circuit and its secondary interconnected with the first rectifier. The thruster power supply may include a d.c. to a.c. switching circuit connected to the solar photovoltaic array, a first rectifier connected to the electric propulsion thruster and a transformer with its primary interconnected with the switching circuit and a first winding of its secondary connected to the first rectifier and a second winding of its secondary connected to a second rectifier connected to the bus loads. The system may further include a battery system interconnected with the second winding of the secondary for charging the battery when the solar array is empowered and supplying power to the primary when the solar array is not empowered. The system may further include a battery system interconnected with the second winding of the secondary for charging when the solar array is empowered and supplying power to the primary when the solar array is not empowered, and supplying power to the bus loads and payloads; the second winding may include a second d.c. to a.c. switching circuit for transferring the battery power through the second winding to the first winding of the secondary to power the electric propulsion thruster. The electric propulsion system may include a Hall effect thruster and the thruster power supply may include a discharge power supply. The electric propulsion thruster may include a number of thrusters each powered by at least one of the thruster power supply sections. At least some of each the thruster power supply sections may be cross-strapped to one or more unassigned thrusters to provide redundant power sources. 
     This invention also features a solar powered spacecraft power system including a solar photovoltaic array, an electric propulsion system connected directly to the solar photovoltaic array in parallel with the spacecraft power system; the electric propulsion system including a Hall effect thruster, a thruster power supply for driving the thruster; a sensor for sensing the power output of the solar array and a controller responsive to the power output of the solar array and configured to periodically adjust an operating parameter of the thruster to operate the thruster at the maximum available output power of the solar array including comparing a previous solar array output power level with a later solar array output power level, and incrementing the operating parameter with a positive value if the later is greater and with a negative value if the later is lesser. 
     In a preferred embodiment the operating parameter may be mass flow rate of the propellant of the Hall thruster. The operating parameter may be the discharge voltage of the Hall thruster. 
    
    
     
       BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS 
       Other objects, features and advantages will occur to those skilled in the art from the following description of a preferred embodiment and the accompanying drawings, in which: 
         FIG. 1  is a schematic block diagram of a conventional, prior art solar powered spacecraft power system; 
         FIG. 2  is a schematic block diagram of a solar powered spacecraft power system for a Hall effect thruster according to one embodiment of this invention; 
         FIG. 3  shows in more detail the thruster power supply and electric propulsion thruster of the electric propulsion system of  FIG. 2 ; 
         FIG. 4  shows another embodiment of a thruster power supply of  FIG. 2 ; 
         FIG. 5  shows another embodiment of a thruster power supply of  FIG. 2 ; 
         FIG. 6  illustrates V-I characteristics and maximum power output point for a solar array; 
         FIG. 7  illustrates the maximum power output characteristics of  FIG. 6  and the V-I characteristics for a thruster at different mass flow conditions; 
         FIG. 8  is a schematic block diagram of an embodiment of an improved solar powered spacecraft power system according to this invention which operates the thruster at maximum available output power of the solar array accommodating for performance degradation; 
         FIG. 9  is a flow diagram of the peak power hunting algorithm configuring the controller, of  FIG. 6 ; and 
         FIG. 10  is a schematic block diagram of a thruster power supply consisting of a number of thruster power supply sections configured in subsets to serve one or a plurality of thrusters and provide redundant back-up. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Aside from the preferred embodiment or embodiments disclosed below, this invention is capable of other embodiments and of being practiced or being carried out in various ways. Thus, it is to be understood that the invention is not limited in its application to the details of construction and the arrangements of components set forth in the following description or illustrated in the drawings. If only one embodiment is described herein, the claims hereof are not to be limited to that embodiment. Moreover, the claims hereof are not to be read restrictively unless there is clear and convincing evidence manifesting a certain exclusion, restriction, or disclaimer. 
     There is shown in  FIG. 1  a conventional solar powered spacecraft power system  10 . Power is generated by a solar photovoltaic array (SP)  12  and flows into the power management and distribution (PMAD) system  14 . From there it is distributed to all the bus loads  16  and payload(s)  18 . In modern spacecraft that use electric propulsion (EP) systems  20 , the EP system  20  constitutes and additional load and can be a major load. A conventional PMAD system  14  may contain a peak power tracking or solar array shunt circuitry. There are a variety of possible PMAD implementations in the prior art. In the more sophisticated form, there is a DC/DC converter that outputs regulated bus voltage that is distributed throughout the bus to the EP system  20 , bus load  16  and payload(s)  18  as well as the battery charger and batteries for energy storage which is used when the spacecraft is in eclipse and the SP delivers no power. When the spacecraft uses an EP system  20  that consumes a substantial portion of the SP  12  output this invention becomes very advantageous. 
     In accordance with one embodiment of this invention,  FIG. 2 , PMAD unit is removed from the power path to EP system  20  and the solar array SP  12  is connected directly to EP system  20 . PMAD unit  14 ′ is now much smaller. It handles much less power and therefore can be smaller resulting in mass and volume savings and because it processes less power it also has lower losses. The heat generated by these power losses in the PMAD must be rejected from the spacecraft to maintain its temperature within operating limits. Heat rejection equipment (radiators) are heavy with a specific mass of the order of 30 kg per kW of rejected heat. This lower loss or equivalently higher efficiency power processing not only results in a smaller array but also in smaller radiators, in both cases significantly reducing the spacecraft mass. EP system  20 ,  FIG. 2 , includes a thruster power supply and a Hall effect thruster. In a Hall effect thruster the thruster power supply is a discharge power supply (DPS). 
     A more detailed diagram of the improved power system of  FIG. 2 , is shown in,  FIG. 3 . Here PMAD unit  14 ′ a  is in parallel with the Hall effect thruster power supply  32 . EP system  20   a  includes a Hall effect thruster  34  and so thruster power supply  32  is implemented with a discharge power supply (DPS). DPS  32  is a DC/DC converter for driving the discharge of Hall effect thruster  34 . Other portions of EP thruster  34 , e.g. electromagnets, cathode heater, which will be discussed in more detail with respect to  FIG. 8 , consume a very small fraction of the power required for EP thruster  34 . Typically DPS  32  processes more than 90% of the power for EP thruster  34 . Thus these auxiliary supplies, e.g. cathode heater, etc. can be easily powered as other bus loads  16   a  with near zero impact on the system efficiency. Typically, bus load  16   a  also includes spacecraft communications, guidance, navigation and control (GNC), command and data handling (C &amp; DH) etc. Thruster power supply, DPS  32  is configured as a DC/DC converter including a.c./d.c. switching circuit  36  and rectifier circuit  38 . Transformer  40  interconnects switching circuit  36  and rectifier  38  switching circuit  36  is connected directly to solar panel  12   a  and to the primary winding  42  of transformer  40 . Rectifier circuit  38  is interconnected with EP thruster  34  and is connected to the secondary winding  44  of transformer  40 . 
     Further improvement can be achieved by combining a portion of PMAD  14 ′ a  with DPS  32  into one DC/DC converter with two secondary windings on its transformer as shown in  FIG. 4 . Here transformer  40   b  includes two secondary windings  44   b  and  44   bb , each of which is connected to a rectifier circuit  38   b  and  38   bb . Rectifier  38   b  directly supplies EP thruster  34   b  so the EP system is directly connected to solar panel  12   b . Second rectifier  38   bb  powered by second winding  44   bb  then drives bus loads  16   b  and payloads  18   b . There is shown a battery system  50  including a battery and a charger circuit interconnected through diodes such as OR-ing diodes  52  and  54  to SP  12   b . When SP  12   b  is providing power some of that power delivered through d.c./a.c. switching circuit  36   b  is transferred through transformer  40   b  to rectifying circuit  38   b  to power EP thruster  34   b . But another portion of that is supplied through winding  44   bb  to second rectifier circuit  38   bb  the output of which powers not only bus loads  16   b  and payloads  18   b  but also charges battery system  50 . Diode  54  blocks battery system  50  from back feeding to solar panel  12   b  and from any noise that may be produced by switching circuit  36   b . When SP  12   b  is producing no power, e.g. it is eclipsed, battery system  50  provides power through diode  52  on lines  56  and  58  back through switching circuit  36   b , and transformer  40   b  to supply rectifier circuit  38   b  which powers EP thruster  34   b  and to supply rectifier  38   bb  which powers bus loads  16   b  and payloads  18   b.    
     An alternative to the use of the OR-ing diodes  52 ,  54  in  FIG. 4 , can be realized by making the DC/DC converter  32  of the DPS bi-directional as shown in  FIG. 5 . Here rectifier circuit  38   cc  includes an addition, a switching circuit  60 . This implementation becomes attractive when the EP system power is comparable to the power consumed by the rest of the spacecraft. When the SP is active power flows via path  1  through from SP  12   c  to switching circuit  36   c  and from there through primary  42   c  of transformer  40   c  to first secondary winding  44   c  through rectifier  38   c  to EP thruster  34   c . Likewise power flows through second winding  44   cc  of the secondary winding through rectifier circuit  38   cc  to supply bus loads  16   c , payloads  18   c  and battery system  50   c . Here, however, when the SP  12   c  is inactive battery power flows back through lines  62  and  64  to power bus loads  16   c  and payloads  18   c  and also through DC/DC converter switching circuit  60  to second winding  40   cc  where it is coupled through the transformer to first winding  44   c  and then through rectifier  38   c  to power thruster  34   c.    
     The thruster  34   b  has a broad operating range and can be used to operate the solar array SP  12  at peak power or off peak as needed. This eliminates the traditional peak power tracker employed in conventional PMAD systems. Connecting directly the main EP power supply or discharge power supply DPS in the case of a Hall thruster has other very important benefits in addition to reducing mass, losses and costs. These benefits are associated with regulation of the array in accordance with this invention as it ages and its output drops. Separate circuitry is normally used in conventional PMAD approaches. This invention, however, offers and elegant solution as described below. The way the EP systems can be used to maintain the SP at its peak power as it ages is illustrated in  FIGS. 6 and 7 .  FIG. 6  shows typical photovoltaic solar array voltage current (V-I) characteristics  80 ,  82  while  FIG. 7  shows typical Hall effect thruster voltage current characteristics with three different mass flows m 1    84 , m 2    86 , and m 3    88 . Also as shown in  FIGS. 6 and 7  are identical constant power curves P 1    90  and P 2    92 . At the beginning of life (BOL) the maximum power point of the solar array,  FIG. 6 , is at a point labeled A. Ignoring losses in the DPS the same power point is shown in the V-I characteristic of the typical Hall effect thruster in  FIG. 7 . The voltage and current values are different at each of the two A points but by energy conservation the power of P 1  in both points As must be the same (again ignoring losses in power processing). To maintain the array at point A, the array output V a  and I a  is measured and the Hall effect thruster DPS voltage setting or the mass flow rate of the propellant through the thrusters are adjusted such that the V a , I a  product is at a maximum. The following example assumes a fixed discharge voltage (V d ) set by the DPS and varying the mass flow. The opposite is an equally valid strategy where the mass is held constant and the voltage V d  is varied or both V d  and mass flow rate of the propellant can be varied subject to power conservation. That is, the operating parameter may be the mass flow rate of the propellant or the discharge or beam voltage. Assume then that initially the thruster is at point B (power equals P 2 , voltage V D , {dot over (m)}={dot over (m)} 2 ). The control algorithm according to one embodiment of this invention that resides in the digital control interface unit (DCIU) described subsequently with respect to  FIGS. 8 and 9  perturbs the mass flow to {dot over (m)} 1 . This increases delivered power to P 1  and shifts the SP and EP to point A. The control algorithm does not know if maximum power was reached and increases the mass flow further to the value {dot over (m)} 3 . However, the array cannot deliver higher power than P 1  and falls off the peak to P 2 , a point labeled C. On the thruster side, the output must also drop to line P 2 . The control algorithm then reduces the mass flow to {dot over (m)} 1  from {dot over (m)} 3  and returns the system to point A. Similar “peak power hunting” process can be carried out at any thruster discharge voltage. Thus using the EP to maintain the array at peak power not only benefits the array and PMAD but also ensures that the thrusters operate at the peak available power at any point in the mission. Because to first order a thruster delivers thrust linearly proportional to its input power maximum power corresponds to maximum thrust which reduces the maneuver time. Alternatively the thruster could be operated at higher specific impulse thus saving fuel. 
     A conventional PMAD which may contain a peak power tracker (that may be continuous or have discrete steps) lowers the voltage as the array ages. However, the EP system must always remain at a power level that is below the temporary maximum power, never at the maximum power to preserve margin for the unknown and unpredictable aging of the array. If one assumes for illustration sake that there are no other loads on the spacecraft power system, the EP operating set point dictates the output power of the array which must be at points B or C,  FIGS. 6 and 7  even at BOL to provide the aforementioned margin. This means that the EP thruster does not get all the power it could get, and hence the thrust is reduced, and transit/maneuver time is extended. With the approach of this invention the EP is used to find the maximum power the array is capable of delivering and then setting the thruster to operate at that power. Thus, the array capability dictates the power to the thruster not the other way around. 
     In a typical EP system, the digital control interface unit (DCIU) sets the thruster operating points (V d  and I d ˜{dot over (m)}) DCIU  100 ,  FIG. 8 , receives digital commands from the spacecraft computer (S/C), and sends out analog set points for the various voltages/currents for each of the converters in  FIG. 7 . DCIU  100  includes typically a microprocessor or microcontroller, memory as well as analog to digital converters and digital to analog converters. The low converters servicing the typical Hall effect thruster include housekeeping DC/DC to converter  102 , cathode heater DC/DC converter  104 , cathode keeper DC/DC converter  106 , magnet A DC/DC converter  108 , magnet B DC/DC converter  110 , valve driver and pressure sensor DC/DC converters  112 , and discharge DC/DC converter  32   d . In  FIG. 8  PMAD  14   d  has associated with it an EMC filter and in rush limiter  114  which is a common approach to prevent backward reflection of converter switching noise and other unwanted signals. Also illustrated in  FIG. 8  is the fact that solar photovoltaic array  12   d  need not be a single solar array but may be formed of a number of solar panels  116 . By far the largest converter is the one associated with the thruster power supply discharge converter  32   d ,  FIG. 8 , which is why in the approach according to this invention the thruster power supply, DPS or discharge converter  32   d  is connected directly to the SP  12   d . The spacecraft computer (S/C) typically will send a command to DCIU  100  to start the thruster and produce thrust for some period of time at a desired operating conditions (e.g. nominal V d , I d ). DCIU  100  executes a series of instructions until the thruster reaches a desired operating point. With this invention, however, DCIU  100  receives an additional command to operate at maximum available power (e.g. to deliver maximum thrust at nominal V d ). DCIU  100  then drives the discharge converter  32   d  to hunt for maximum power from SP  12   d  or specific solar panels  116 . In one implementation DCIU  100  may receive power readings directly from a watt-meter or volt times amp meter  120 ,  FIG. 8 , and calculate the set point for each DPS converter. This requires an additional algorithm in DCIU  100 . Alternatively, each DPS converter may do its own processing to draw maximum power from the array segment. DCIU  100  then sums the output current from all DPS converters and alters the mass flow set point (valve position see converter  112 ) for the thruster that is fed by the group of converters. Depending upon the health of the various sections of solar panel, the individual converters may deliver substantially different power to the thruster while each section delivers the maximum it can. 
     A maximum power algorithm  130 ,  FIG. 9 , according to one embodiment of this invention causes DCUI  100  to monitor the power at meter  120  and stores that value  132 . The present sensed value P i+1  is compared to the previous power P i  at  134 . If the present power reading P i+1  is greater than the previous power reading P i  then the system is incremented to a new mass flow {dot over (m)} i+1  by adding a mass Δm to the previous mass {dot over (m)} i    136 . If P i+1  is not greater than P i  then  138  the Δm increment is applied as a negative, that is −Δm. The instruction is sent on line  140  to the flow control valve driver  112  and the absolute value of Δm is forwarded  142 . The loop speed for this adjustment need not be extremely high, adjustment even every few minutes may be found sufficient. 
     Note that the maximum power algorithm,  FIG. 9 , is uninformed of the power processed by the DPS  32   e  or the power processed by the PMAD  14 ′, the former being consumed by the hall thruster plasma discharge and the latter by the bus loads plus the payload. The only requirement is that their sum must be equal to the product Ia*Va ( 120 ), again ignoring losses in the power processing. The SP,  12   e , power output (=Ia*Va) can be changing and the bus load and the payload,  18   e , can also be changing while the algorithm automatically accommodates these simultaneous changes by directing the excess power to the DPS,  32   e , always giving priority to the bus and payload power needs. This can be explained by the following example. Suppose that that the EP is off (consumes no power) and the bus plus the payload consumes power P 2  as shown in  FIG. 6 . At BOL, the operating point of the array then must be at either point B or C. Then the EP thruster is commanded to start at first at very low power which it gradually increases using the algorithm in  FIG. 9 . This will be shifting the power produced by the array toward P 1  and point A. If suddenly, the bus load demands more power, the SP cannot supply it and the algorithm in  FIG. 9  reduces the SP power by reducing the power to the EP thruster. In this manner, the EP thruster continually drives the SP to maximum power point while automatically satisfying the power demand by the bus load and the payload. 
     So far, simple block diagrams have been used to describe the invention but large high power spacecraft can be much more complicated. For example as shown in  FIG. 10  the improved power system of this invention for a high power spacecraft may have 30 kW of power on board. The solar array  12   f  may include ten solar panels  116   f . Each panel is connected to a discharge power supply DC converter  32   f . The outputs of five of the converters are paralleled and feed 15 kW Hall thruster  34   f . The other five converters  32   f  deliver power to a second Hall thruster  34   f . If needed the outputs can be cross-strapped such as by switches  150  such that either thruster  34   f  can be connected to either set of the five converters  32   f  and their respective arrays  116   f . This is desirable for redundancy. Power for the rest of the bus and payloads may be delivered through a separate low power DC/DC converter that outputs the conventional 28 volts. This converter can be fed from one or more of the solar panels  116   f  or may have its own solar panel in the solar array. 
     Although specific features of the invention are shown in some drawings and not in others, this is for convenience only as each feature may be combined with any or all of the other features in accordance with the invention. The words “including”, “comprising”, “having”, and “with” as used herein are to be interpreted broadly and comprehensively and are not limited to any physical interconnection. Moreover, any embodiments disclosed in the subject application are not to be taken as the only possible embodiments. 
     In addition, any amendment presented during the prosecution of the patent application for this patent is not a disclaimer of any claim element presented in the application as filed: those skilled in the art cannot reasonably be expected to draft a claim that would literally encompass all possible equivalents, many equivalents will be unforeseeable at the time of the amendment and are beyond a fair interpretation of what is to be surrendered (if anything), the rationale underlying the amendment may bear no more than a tangential relation to many equivalents, and/or there are many other reasons the applicant can not be expected to describe certain insubstantial substitutes for any claim element amended. 
     Other embodiments will occur to those skilled in the art and are within the following claims.