Abstract:
A system for single wire secondary distribution comprising a spacecraft platform; a central bus interface unit coupled to the spacecraft platform; a payload unit coupled to the central bus interface unit; and a centralized power supply for powering the central bus interface unit and the payload unit; wherein the spacecraft platform provides a command to the central bus interface unit; wherein the central bus interface unit interrupts the power to the payload unit in a manner corresponding to the commands received by the central bus interface unit; wherein the payload unit decodes the interruption to the power and executes the command from the spacecraft platform.

Description:
BACKGROUND OF INVENTION  
       [0001]     1. Field of the Invention  
         [0002]     The present invention relates generally to the transfer of data. More specifically, but without limitation thereto, the present invention relates to a centralized bus interface for connecting a payload unit and a spacecraft bus using a single wire secondary distribution.  
         [0003]     2. Discussion of the Background Art  
         [0004]     Currently the interface between a spacecraft bus and a plurality of payload units requires a separate power line, communication lines and telemetry lines for every payload unit on the spacecraft. The spacecraft bus is the main bus on the spacecraft that is connected to the spacecraft command processor. The spacecraft command processor is the main processor on a spacecraft that sends and receives data from the ground and also sends and receives data from the plurality of payload units on the spacecraft. The plurality of payload units are the electronic sub-systems such as receivers, power amplifiers, frequency converters and low noise amplifiers.  
         [0005]     Prior designs include 8 or more wires connecting one payload unit to the spacecraft bus. Each payload unit on the spacecraft has a separate 8 wire connection to the spacecraft bus including 2 wires for power, 4 wires for commands, 2 wires for telemetry and return wires via the chassis. As the size of payload electronics shrinks, the size of the circuitry interfacing with the spacecraft bus must also shrink in order to properly operate. One current approach to reducing the size and weight of the interface circuitry to each payload unit is utilizing smaller components and higher density packaging technologies. Although this has resulted in very good incremental improvements, a new architecture for these interfaces is needed to make a step decrease in the size and weight of the payload units as each payload unit still requires at least an 8 wire interface.  
         [0006]     Thus there is a need for an interface design between a spacecraft bus and the payload units which can solve the problems discussed above.  
       SUMMARY OF INVENTION  
       [0007]     The present invention advantageously addresses the need above as well as other needs by providing an interface between the payload units and the spacecraft bus using a single wire secondary distribution.  
         [0008]     In one embodiment, the present invention can be characterized as a method of transmitting data to a payload unit comprising the steps of sending a command from a command processor to a central bus interface unit for routing the command to the payload unit; providing power over a power line from a central power supply to the payload unit; interrupting the power to the payload unit to provide the payload unit with the command from the command processor.  
         [0009]     In another embodiment, the invention can be characterized as a method of communicating with a payload unit comprising the steps of providing power to the payload unit over a wire; and providing telemetry from the payload unit over the wire to a spacecraft command processor.  
         [0010]     In yet another embodiment, the invention can be characterized as a system comprising a spacecraft command processor; a central bus interface unit coupled to the command processor; a payload unit coupled to the central bus interface unit; and a centralized power supply for powering the central bus interface unit and the payload unit; wherein the spacecraft command processor provides a command to the central bus interface unit; wherein the central bus interface unit interrupts the power to the payload unit in a manner corresponding to the commands received by the central bus interface unit; wherein the payload unit decodes the interrupts of the power and executes the command from the spacecraft command processor.  
         [0011]     In an additional embodiment the present invention can be characterized as a system for communicating with a payload unit comprising a command processor; and the payload unit coupled to the command processor through a combined power and communication wire; wherein the payload unit receives power and command data over the combined power and communication wire.  
         [0012]     In a further embodiment, the present invention can be characterized as a system for connecting a spacecraft bus to a payload unit comprising an interface for directing a command from a spacecraft command processor; a central bus interface unit coupled to the interface, the central bus interface unit comprising a command decoder; a register coupled to the command decoder; and a switch coupled to the register; wherein the register operates the switch; wherein the switch interrupts an output voltage, the interruption of the output voltage corresponding to the command from the spacecraft command processor; a centralized power supply coupled to the central bus interface unit; and an end user interface coupled to the output voltage, the end user interface comprising a decoder coupled to the output voltage for decoding the interrupts; and a power voltage for powering the payload unit during the interrupts. 
     
    
     BRIEF DESCRIPTION OF DRAWINGS  
       [0013]     The present invention is illustrated by way of example and not limitation in the accompanying figures, in which like references indicate similar elements, and in which:  
         [0014]      FIG. 1  is a block diagram illustrating an overview of the present invention;  
         [0015]      FIG. 2  is a block diagram illustrating a centralized bus interface unit with a single wire power, telemetry, and command distribution in accordance with the present invention;  
         [0016]      FIG. 3  is a schematic diagram illustrating an internally redundant power and command distribution circuitry in accordance with the embodiment shown in  FIG. 2 ; and  
         [0017]      FIG. 4  is a schematic diagram illustrating an interface for powering a payload unit and decoding commands in accordance with the embodiment shown in  FIG. 2 . 
     
    
       [0018]     Skilled artisans will appreciate that elements in the figures are illustrated for simplicity and clarity and have not necessarily been drawn to scale. For example, the dimensions of some of the elements in the figures may be exaggerated relative to other elements to help to improve understanding of embodiments of the present invention.  
       DETAILED DESCRIPTION  
       [0019]     Advantageously, the embodiments described herein below include an apparatus and method for dramatically reducing the size, weight and cost of the electronics needed to provide the interface between a payload unit and a spacecraft bus or platform with which it interfaces. The spacecraft platform provides power and commands to units on the spacecraft and processes telemetry signals.  
         [0020]     Typical prior designs have at least 8 wires for each payload unit connected to the spacecraft bus: 2 wires for power, 4 wires for commands, 2 wires for telemetry and return wires via the chassis. The spacecraft platform interfaces with the ground station and provides all housekeeping functions for the payload which comprises a plurality of payload units. The payload units, e.g., receivers, frequency converters, power amplifiers, low noise amplifiers, send telemetry back to the spacecraft platform and ground station to allow assessment of the health and condition of the payload. In prior designs each payload unit is connected to the spacecraft bus through at least 8 wires.  
         [0021]     Advantageously, the present invention reduces the number of wires from each payload unit to the spacecraft bus to 1 with an additional 10 to 12 wires connecting the spacecraft bus to a central bus interface unit. Thus, in the present embodiment, the spacecraft bus is connected to a central bus interface unit which is in turn connected to the plurality of payload unit through only 1 wire per payload unit. Thus, for a design with 20 payload units connected to the spacecraft bus, the total number of wires is reduced from 160 to 32, an 80% decrease.  
         [0022]     Additionally, the present embodiment is much more reliable than the prior designs because the number of wires connecting the spacecraft bus to the payload units is greatly reduced. Each one of these wires has the possibility of failing or introducing errors into the system, thus the reduction in wires reduces the chances that there will be a failure. By decreasing the wires, e.g., by 80% as described above, the present embodiment can also be assembled and tested in much less time than prior designs because there are less connections between the payload units and the spacecraft bus to test.  
         [0023]     Referring to  FIG. 1 a  block diagram is shown illustrating an overview of the present invention. Shown is a spacecraft platform  50 , a spacecraft bus interface  60 , a central bus interface unit  108 , and a plurality of outputs  110  to a plurality of payload units.  
         [0024]     The spacecraft platform  50  is connected via the spacecraft bus interface  60  to the central bus interface unit  108 . The central bus interface unit has a plurality of outputs  110 .  
         [0025]     The plurality of outputs  110  are connected to a plurality of payload units. Each payload unit corresponds to one of the plurality of outputs  110 . In accordance with the present invention the plurality of outputs  110  are all single wire distributions for providing both power and command data over a single wire.  
         [0026]     Prior designs did not incorporate the central bus interface unit. The prior designs had each payload unit connected to the spacecraft bus with at least 8 wires, each of the wires used separately for power, command data and telemetry. Thus, the present invention greatly reduces the number of wires needed for the interface between a plurality of payload units and the spacecraft bus  60  by providing a single wire to each payload unit that is used for both power and command data.  
         [0027]     Referring to  FIG. 2 a  block diagram is shown illustrating a centralized bus interface unit with a single wire power, telemetry, and command distribution in accordance with one embodiment of the present invention. Shown is a plurality of centralized power supplies  102 , a power OR block  104 , a central bus interface unit  108 , a spacecraft interface  112 , a main power supply  100 , a central power supply line  106 , a plurality of outputs  110 , a command bus  118 , a telemetry bus  120 , an data bus  114 , a plurality of power supply control lines  116 , and a plurality of payload units  122 .  
         [0028]     The main power supply bus  100 , command bus  118  and the telemetry bus  120  are all part of the spacecraft bus interface  60  which is connected to the spacecraft platform  50 , shown in  FIG. 1 . The main power supply  100  is connected to the plurality of centralized power supplies  102 . The command bus  118  and the telemetry bus  120  are connected to the spacecraft interface  112 . The spacecraft interface  112  is connected to the plurality of centralized power supplies through a plurality of power supply control lines  116 . The spacecraft interface  112  is also connected to the central bus interface unit  108  through the data bus  114 .  
         [0029]     The plurality of centralized power supplies  102  are connected to the power OR block  104  which is connected to the central bus interface unit  108  through the central power supply line  106 . The central bus interface unit  108  is connected to the plurality of payload units  122  through the plurality of outputs  110 .  
         [0030]     As is shown by the overview in  FIG. 1 , the plurality of centralized power supplies  102 , the power OR block  104 , and the spacecraft interface  112  are all optional features of the present invention. The main power supply  100 , the command bus  118  and the telemetry bus  120  can be connected directly to the central bus interface unit  108 . In this embodiment some of the function of the spacecraft interface may be included in the central bus interface unit  108 . The plurality of centralized power supplies  102 , the power OF block  104 , and the spacecraft interface are optional features, however, in a preferred embodiment they are included because they add redundancy and reliability as compared to having only one power supply. This is because if one of the plurality of centralized power supplies  102  fails, the power OR block  104  will still provide power to the central bus interface unit  108 .  
         [0031]     The main power supply bus  100  from the spacecraft platform is connected to the plurality of centralized power supplies  102 . The plurality of centralized power supplies  102  are a plurality of power supply slices that are connected in parallel and provide redundancy if one or more of the plurality of centralized power supplies  102  fails. The outputs from the plurality of centralized power supplies  102  are connected to a power OR block  104 . The power OR block  104 , is a logic device that supplies a high output as long as one of the plurality of centralized power supplies is outputting power into the power OR block  104 . The output from the power OR block  104  is the central power supply line  106 . The central power supply line  106  provides power to the central bus interface unit  108 . The central bus interface unit  108  is later described in more detail with reference to  FIG. 3 . In an alternative design, the main power supply  100  from the spacecraft bus can be directly connected to the central bus interface unit  108 .  
         [0032]     The spacecraft platform  50  provides commands through the command bus  118  which is part of the spacecraft bus interface  60 . The commands are sent over the command bus  118  through the spacecraft interface  112 . The commands are then directed either to the plurality of centralized power supplies  102  or to the central bus interface unit  108 . In one embodiment, the commands include a command header, which contains information as to where the command is to be directed. Each of the plurality of centralized power supplies  102  and each of the plurality of payload units  122  has a unique command header associated with it. Thus, when the spacecraft platform  50  sends out a command, the spacecraft interface  112  will direct the command accordingly. The spacecraft interface  112  will decide, based upon the command header, if the command is for one of the plurality of centralized power supplies  102 , e.g., an on/off command, or if the command is for one of the plurality of payload units  122 .  
         [0033]     If the command is for one of the plurality of centralized power supplies  102  the command will be sent to one of the plurality of centralized power supplies  102  through one of the plurality of power supply control lines  116 . For example, the spacecraft platform  50  can individually turn on or off each of the plurality of centralized power supplies  102  by sending an on/off command over the command bus  118  to the spacecraft interface  112 . The spacecraft interface  112  then sends the on/off command over one of the plurality of power supply control lines  116  to one of the plurality of centralized power supplies  102 .  
         [0034]     If, however, the command, including the command header, from the spacecraft platform  50  is for one of the plurality of payload units  122  the command will be directed from the spacecraft interface  112  to the central bus interface unit  108  through the data bus  114 . The spacecraft interface  112  will direct the command to the central bus interface unit  108  based upon the command header. The central bus interface unit  108  interprets the command header and directs the command over one of the plurality of outputs  110  to one of the plurality of payload units  122 .  
         [0035]     The plurality of outputs  110  provide both power and data, i.e., the commands from the spacecraft platform  50 , to a plurality of payload units  122 . This will be described in greater detail herein with reference to  FIGS. 3 and 4 . The plurality of output  110  are each a single wire. Each payload unit  122  is provided with power and data over one of the plurality of outputs  110 , thus, greatly reducing the number of wires required from the spacecraft bus to the plurality of payload units  122  as compared with the prior designs described above. Telemetry is also provided from the plurality of payload units  122  back to the central bus interface unit  108  and back to the spacecraft platform  50 . This will be described in greater detail herein with reference to  FIG. 3 . In a typical application, the plurality of payload units  122  can range from about 20 payload units to 150 payload units, however, the number of payload units  122  does not limit the present invention, for example, in one variation only one payload unit is connected to the central bus interface unit  108 .  
         [0036]     Referring to  FIG. 3 a  schematic diagram is shown illustrating one embodiment of the central bus interface unit  108  shown in  FIG. 2 . Shown is the central power supply line  106 , a fault isolation circuit  200 , a voltage regulation circuit  202 , an output voltage  212 , a decoder  208 , a register  210 , a bias voltage  206 , a op amp  220 , a reference voltage  214 , a Vcc voltage  216 , a first transistor  250 , a second transistor  252 , a third transistor  254 , a fourth transistor  256 , and a telemetry line  222 .  
         [0037]     The central power supply line  106  is the output from the power OR block  104  shown in  FIG. 2 . The central power supply line is connected to the drain of the third transistor  254 . The third transistor  254  is part of to the fault isolation circuit  200 . The source of the third transistor  254  is connected to the drain of the fourth transistor  256 . The source of the fourth transistor  256  is the output voltage  212 .  
         [0038]     The command decoder  208  is connected to the register  210 . The output of the register is coupled through an impedance to the gate of the first transistor  250  and to the gate of the second transistor  252 . The drain of the first transistor  250  is coupled through an impedance to the gate of the third transistor  254 . The drain of the second transistor  252  is coupled through an impedance to the gate of the fourth transistor  256 . The sources of both the first transistor  250  and the second transistor  252  are connected to ground.  
         [0039]      FIG. 3  represents part of the central bus interface unit  108  of  FIG. 2 . The central bus interface unit  108  generally includes more than one of the circuits shown in  FIG. 2 . Specifically, in one embodiment, the central bus interface unit  108  will include the circuitry shown in  FIG. 2  for each payload unit  122  that is connected to the central bus interface unit  108 . The payload unit  122  can be many types of circuits, including for example, a receiver, a power amplifier, a frequency converter, and a low noise amplifier. The present embodiment is not limited to particular types of payload units  122 , but rather, as will be appreciated by one of ordinary skill in the art, many different types of payload units  122  may be used.  
         [0040]     The decoder  208  receives commands from the spacecraft platform  50  over the data bus  114 . The decoder  208  is connected to the register  210 . The commands from spacecraft platform  50  are translated into short pulses by the decoder  208  and the register  210 . When the register  210  outputs a high signal, the output voltage to the payload unit  122  is momentarily interrupted (e.g., for 10 mSec). This interruption is received by a decoder at the payload unit  122  and thus receives the commands from the spacecraft platform. The decoder is described in greater detail herein with reference to  FIG. 3 .  
         [0041]     In operation, when, the register  210  outputs a high signal in order to momentarily interrupt the output voltage  212 , this applies a high voltage to the gate of the first transistor  250  and the gate of the second transistor  252 , thus turning on the first transistor  250  and the second transistor  252 . This brings the bias voltage at the gate of the third transistor  254  and the fourth transistor  256  below the level needed to turn on the third transistor  254  and the fourth transistor  256 . Thus, the third transistor  254  and the fourth transistor  256  turn off. This causes the momentary interruption of the output voltage  212  until the register  210  outputs a low signal. When the register outputs a low signal, the voltage at the gate of the third transistor  254  and the fourth transistor  256  returns to a high enough level to turn back on the third transistor  254  and the fourth transistor  256 . This restores the output voltage  212 . The biasing levels and operation of transistors is known to one of ordinary skill in the art.  
         [0042]     Generally, the output voltage to the payload unit  122  will be interrupted multiple times for each command from the spacecraft command processor. As shown, the register  210  is redundant in case one of the registers fails the power will still be interrupted. Additionally, either of the fault isolation circuit  200  or the voltage regulation circuit  202  can interrupt the output voltage  212  such that the command is provided to the payload unit  122 . The third transistor  254  and the fourth transistor  256  act as switches and are located inside the fault isolation circuit  200  and the voltage regulation circuit  202  respectively. The switches inside of the fault isolation circuit  200  and the voltage regulation circuit  202  are connected in series such that if one of the switches fails, the other switch will still be able to interrupt the output voltage  212 .  
         [0043]     In addition to transferring data sent by the spacecraft command processor  50 , the circuitry shown in  FIG. 3  also provides redundancy and short circuit current protection. Specifically, the fault isolation circuit  200  also senses the current to each payload unit  122  and disconnects if an over current fault condition is detected, thus isolating a short circuit in one payload unit  122  from affecting other payload units  122 . The voltage regulation circuit also has the additional functionality of post regulation, providing each payload unit  122  with a cleaner additionally refined power form.  
         [0044]     The telemetry line  222  is coupled to a demodulator circuit (not shown) in the central bus interface unit  108 . The demodulator circuit then receives the telemetry data received from the payload unit  122  and sends it back to the spacecraft platform through the data bus  114  and interface circuit  112 . The telemetry functions can be added to the decoder chip  350  of  FIG. 4  or can be implemented on a separate ASIC at the payload unit  122 . A variety of techniques can be used for transmitting telemetry data from the payload unit  122  back to the spacecraft command processor. One technique is to use a spread spectrum signal generated by the decoder chip  350  back to the demodulator (not shown) which can be added to either the fault isolation circuit  200  or the voltage regulation circuit  202 . Notch filters can be used on the power line in each unit as well as in the central bus interface unit  108  to prevent the spread spectrum signal from propagating to undesirable locations through the various power lines. The single wire from the central bus interface unit  108  to each payload unit  122  is generally a shielded wire or coax cable to control radiated emissions and susceptibility to outside signals, however, any wire or transmission medium can be used in applications which do not require shielding. In an alternative form the telemetry data can be sent over a single frequency modulated carrier or any other of a variety of know transmission methods. Additionally, the same type of power interruption that is used to transmit the command data can be used in transmitting the telemetry data back to the central bus interface unit  108 .  
         [0045]     In an alternative embodiment, the telemetry data can be transmitted back to the central bus interface unit  108  using a separate return wire. Additionally, if more than one wire is used in between the central bus interface unit  108  and the payload unit  108  in accordance with the present invention, power can be supplied either on the line where telemetry data is being sent or on the same line the command data is being sent. Thus, in accordance with the present invention, power is supplied over the same line as either the command data or telemetry data or both.  
         [0046]     The bias voltage  206  can be supplied directly from either a separate output of the power OR block  104  or from the central power supply line  106 . Alternatively, the bias voltage  206  can be connected to the central power supply  106  through a diode and capacitor, thus providing a bias voltage  206  which is lower than the central power supply  106 .  
         [0047]     In another embodiment, both the commands and telemetry are transmitted on the power line using a spread spectrum signal. Each payload unit  122  and the central bus interface unit  108  can include a spread spectrum demodulator and modulator, respectively. In this embodiment, the power is not interrupted but the spread spectrum signal is sent to the payload unit  122  on the same line that power is being supplied to the payload unit  122 . In this embodiment, the central bus interface unit  108  receives a command from the spacecraft platform  50 . The central bus interface unit then sends the command to one of the payload units  122  using a spread spectrum signal. As described above, the command includes a command header which identifies which of the plurality of payload units  122  should receive the command. The central bus interface unit  108  includes a spread spectrum modulator which enable it to send out the signal to the desired payload unit  122 . In this embodiment the payload unit  122  includes a spread spectrum demodulator so that it can receive the command. The payload unit  122  receives the command over the power line. Thus, the output power  212  from the central bus interface unit  108  is also a communication line, i.e., a combined power and communication line.  
         [0048]     In an alternative embodiment, the commands can be transmitted about a 10 Mhz frequency modulated carrier and the telemetry data can be transmitted about a 15 Mhz frequency modulated carrier on the same line as the power to the payload unit  122 . The central bus interface unit  108  includes a frequency modulator. The central bus interface unit  108  receives the commands from the spacecraft platform  50 . The commands are then modulated and sent to the payload unit  122 . The payload unit  122  includes a frequency demodulator, such that the commands being sent to the payload unit  122  on the 10 Mhz frequency modulated carrier can be demodulated. This allows the payload unit to properly receive the commands from the spacecraft platform  50 . In this embodiment, the power and data are still transmitted over the output power line  212 .  
         [0049]     Referring to  FIG. 4 a  schematic diagram is shown illustrating an interface for powering a payload unit and decoding commands in accordance with the embodiment shown in  FIG. 2 . Shown is a decoder chip  350 , an input voltage  300 , a power voltage  302 , a power reset  304 , a pulse width decoder  306 , a counter  308 , a reset line  310 , an output  312 , a diode  314  and a capacitor  316 .  
         [0050]     The input voltage  300  is connected to the diode  314 . The diode is connected to a the capacitor  316  and the capacitor  316  is connected to ground. The node between the diode  314  and the capacitor is the power voltage  302 . The power voltage  302  is connected to the pulse width decoder  306 , the counter  308 , and the power reset  304 . The reset line  310  of the power reset  304  is connected to the counter  308 . The input of the pulse width decoder  306  is connected to the input voltage  300  and the output of the pulse width decoder  306  is connected to the counter  308 . The pulse width decoder  306 , the power reset  304  and the counter  308  comprises one embodiment of the decoder chip  350 .  
         [0051]     The interface includes circuitry for powering the payload unit  122  and for decoding the commands sent by the central bus interface unit  108 . As described above, in one preferred embodiment, the commands are sent by momentarily interrupting the output power  212 .  
         [0052]     In a preferred embodiment, each payload unit  122  will have the interface as shown in  FIG. 4 . This allows each payload unit  122  to be able to process commands from the spacecraft command processor. The output voltage  212  of  FIG. 3  is connected to the input voltage  300  of  FIG. 4 . By interrupting the input voltage to the payload unit  122 , commands are sent to each payload unit  122  over a single wire.  
         [0053]     The power voltage  302  supplies the payload unit  122  with a constant voltage such that when the input voltage  300  is momentarily interrupted the payload unit will not turn off. The decoder  306  detects the momentary interruptions of the input voltage  300  and sends a valid pulse signal to the counter  308 . The counter  308  sends an output  312  which corresponds to the original command sent by the spacecraft platform  50  to the payload unit  122 . As described earlier, each command from the spacecraft platform will generally translate into a plurality of pulse signals or interruptions of the power to the payload unit  122 . The payload unit  122  can then respond to the commands from the spacecraft platform.  
         [0054]     The power reset  304  will reset the payload unit if the power to the payload unit is interrupted for a long enough period of time, e.g., a time longer than the interrupt time for a valid pulse. The power reset  304  sends a reset signal to the counter which is relayed to the payload unit  122  through the counter  308 . This will reset the entire payload unit  122 .  
         [0055]     Alternative to the interface shown in  FIG. 4 , the interface can include either a spread spectrum demodulator or a frequency demodulator. In this embodiment, the modulator chosen corresponds to the type of signal being generated by the central bus interface unit  108 . The output from either demodulator is provided to the payload unit  108 .  
         [0056]     The present invention has been described herein as being used for application on a spacecraft, however, the present invention can also be used in any application where the cost of wire is at a premium or where efficiency and reliability are desired, e.g., on an airplane.  
         [0057]     The present invention as described herein includes many redundant circuit parts, e.g., the plurality of centralized power supplies  102 . For application in a space environment, many redundancy features can be very important as repairing a part that has failed in space is either very costly or impossible. However, in accordance with the present invention, this redundancy is not required in all applications. This is because in some applications, repair is possible, or a very low cost design is necessary. Every time you add a redundancy it will add cost to the apparatus.  
         [0058]     For example, one type of redundancy that is not required in some embodiments is the plurality of centralized power supplies  102 . In this embodiment, the central bus interface unit is powered directly from the main power supply line  100  of the spacecraft bus  60 . Another example of this redundancy is having both the fault isolation circuit  200  and the voltage regulation circuit  202  being able to interrupt the central supply voltage  204 . If either the fault isolation circuit  200  or the voltage regulation circuit  202  fails, the power to the payload unit  122  will still be interrupted. In contrast, in another embodiment, only one of the fault isolation circuit  200  and the voltage regulation circuit  202  is able to interrupt the power to the payload unit  122  when sending a command to the payload unit  122 .  
         [0059]     While the invention herein disclosed has been described by means of specific embodiments and applications thereof, other modifications, variations, and arrangements of the present invention may be made in accordance with the above teachings other than as specifically described to practice the invention within the spirit and scope defined by the following claims.