Patent Publication Number: US-9413162-B2

Title: Modular equipment center distributed independent protections

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     The present U.S. Nonprovisional application is related to co-pending U.S. Nonprovisional application Ser. No. entitled “MODULAR EQUIPMENT CENTER DISTRIBUTED PRIMARY POWER ARCHITECTURE” Ser. No. 14/052,327, U.S. Nonprovisional application Ser. No. entitled “MODULAR EQUIPMENT CENTER ZONAL STANDALONE POWER SYSTEM CONTROL ARCHITECTURE” Ser. No. 14/052,396, U.S. Nonprovisional application Ser. No. entitled “MODULAR EQUIPMENT CENTER SOLID STATE PRIMARY POWER SWITCHING NETWORK” Ser. No. 14/052,426, U.S. Nonprovisional application Ser. No. entitled “MODULAR EQUIPMENT CENTER DISTRIBUTED EQUIPMENT PACKAGING TRUSS” Ser. No. 14/052,304, U.S. Nonprovisional application Ser. No. entitled “MODULAR EQUIPMENT CENTER INTEGRATED TRUSS SENSORS” Ser. No. 14/052,279, U.S. Nonprovisional application Ser. No. entitled “MODULAR EQUIPMENT CENTER LIGHTNING THREAT REDUCTION ARCHITECTURE” Ser. No. 14/052,292, and U.S. Nonprovisional application Ser. No. entitled “REMOTE MODULAR EQUIPMENT CENTER ARCHITECTURE” Ser. No. 14/052,450, which are all incorporated herein by reference in their entirety, having been filed concurrently with the present application. 
     TECHNICAL FIELD 
     The field of the embodiments presented herein is directed toward modular vehicle architectures, and more particularly, to electronic protections within distributed power and data aircraft architectures. 
     BACKGROUND 
     Most commercial aircraft have one or more centralized equipment bays for housing electrical power and communications equipment. Power and data are distributed from the centralized equipment bays throughout the entire aircraft to control all functions within the aircraft. The centralized equipment bays are displaced from one another across one or more section breaks in the aircraft. Typically, one centralized equipment bay is in a forward section and the other is in an aft section of the aircraft. 
     Generators driven by the main propulsive engines generate three-phase primary electrical power for the aircraft. The primary power is first routed to the aft equipment bay and then through the aircraft to the forward equipment bay. The primary power is then centrally configured for distribution throughout the rest of the aircraft to service various equipment loads. Centralized bus power control units within the equipment bays control all power functions throughout the aircraft. After the centralized conversions, secondary power is routed to remote power distribution units to service the equipment loads throughout the aircraft or directly to equipment loads. 
     All functions of the aircraft are reliant upon the centralized power and communications equipment. If either the power or data from the centralized equipment bays is severed, the receiving equipment goes into a standby state where it becomes difficult for the flight crew to determine the state of the corresponding systems. Also, the backbone of the communication network must be oversized because of the high bandwidth demands during peak times to and from the centralized communication equipment. 
     Composite aircraft do not have an aluminum chassis to serve as the return current path or network. Consequently, either a complex network of wires must be added to provide a current return path for all circuits or dedicated return wires must be added for each equipment load. For example, conductive wiring must be added that extend longitudinally along the length of the composite aircraft as well as laterally across the width of the composite aircraft, as described in U.S. Pat. No. 8,031,458 entitled CURRENT RETURN NETWORK, and which is herein incorporated by reference in its entirety. This solution adds cost, manufacturing and maintenance complexity, increased voltage drop, and undesirable weight to the composite aircraft. Thus, attempts to reduce weight in composite aircraft by minimizing wiring have been counteracted by the need for increased lightning protection components and other reasons in composite aircraft. 
     The aluminum chassis (e.g. components that make up the frame or skin or combination thereof) of traditional aircraft, as well as any other conductive metal structure of the aircraft, is tied together to form a current return network for returning a voltage reference point to the source distribution grounding point. The current return network also provides lightning protections as well as personal safety protection path. However, in composite aircraft where the chassis may be formed of an insulation material, the routing of wires from the generators, to the forward and aft equipment bays, to the remote power distribution units and the equipment loads they service, and back to the forward equipment bay via the current return network, creates a large wire loop. In a composite aircraft, this long wire loop may induce a large current during a lighting strike to the aircraft under certain conditions. To address this concern, the wire loop may be shielded but this large wire loop and its shielding would undesirably contribute a significant amount of weight in the aircraft. 
     Commercial aircraft may be manufactured in separate sections that are then connected together to assemble a complete aircraft. Various systems in the aircraft may have components that are distributed across multiple sections. Before the sections are finally assembled together, many of the components in a section are installed and tested to confirm that they were assembled correctly. Therefore, to test and verify a section, the portions of the systems that are not yet present in the build sequence have to be emulated. Once section installations have been tested, final assembly of the sections forming the aircraft can be performed that would make repairs to errors found after this stage more difficult to correct due to limited accessibility. 
     In today&#39;s aircraft, one of the reasons final assembly is such a time consuming process is because of the large number of primary and secondary power connections and the large number of data connections between adjacent sections. Aircraft could be built at a faster rate and orders for completed aircraft could be filled more quickly by functionally testing systems earlier in the build cycle, thus eliminating the need to emulate some equipment located in other parts of the aircraft, reducing the number of connections across section breaks, eliminating integration panels, and by minimizing the weight and complexity of aircraft wiring. 
     It is with respect to these and other considerations that the disclosure herein is presented. 
     SUMMARY 
     It should be appreciated that this Summary is provided to introduce a selection of concepts in a simplified form that are further described below in the Detailed Description. This Summary is not intended to be used to limit the scope of the claimed subject matter. 
     According to one embodiment disclosed herein, a system for providing distributed electrical protections for electrical systems in a vehicle is provided. The system comprises modular equipment centers (MECs) spatially distributed throughout the vehicle. The system further comprises power controllers within each MEC for implementing protective functions by monitoring and controlling the electrical systems of the aircraft. One or more of the power controllers is configured to implement localized protective functions and perform independent fault assessments. One or more other power controllers within the vehicle are configured to perform coordinated fault assessments. Multiple power controllers within the vehicle may implement multiple protective actions substantially simultaneously without communication from other MECs. 
     According to another embodiment disclosed herein, a method for providing distributed protection and control architecture for electrical systems of a vehicle is provided. The method comprises spatially distributing MECs throughout the vehicle, distributing primary power from a power source to each of the MECs, distributing secondary power from each MEC to equipment loads nearest each MEC, providing integrated protection chipsets (IPCs) where each IPC provides electronic protective functions, performing high power level protections where primary power is received at each MEC from the power source, and independently performing low level protections where secondary power is distributed from each MEC to the equipment loads associated with each MEC. 
     According to yet another embodiment disclosed herein, a system for providing distributed protection of electrical systems in a vehicle is provided. The system comprises MECs spatially distributed throughout the vehicle. One or more power sources provide primary power to the MECs and one or more equipment loads are serviced by each of the MEC. The system further comprises IPCs for monitoring and controlling electrical systems within the aircraft. The IPCs are configured to implement a plurality of protective functions and one or more of the IPC are configured to make independent fault assessments. 
     The features, functions, and advantages that have been discussed can be achieved independently in various embodiments of the present disclosure or may be combined in yet other embodiments, further details of which can be seen with reference to the following description and drawings. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The embodiments presented herein will become more fully understood from the detailed description and the accompanying drawings, wherein: 
         FIG. 1  illustrates a top view of one configuration of an aircraft with spatially distributed modular equipment centers (MECs) wherein equipment loads are serviced by the nearest MEC according to at least one embodiment disclosed herein, 
         FIG. 2  illustrates the splitting of two generators per aircraft engine relative forward and aft of the aircraft according to at least one embodiment disclosed herein, 
         FIG. 3  illustrates one configuration of primary power feeders connected to generators energizing a power bus network according to at least one embodiment disclosed herein, 
         FIG. 4  illustrates one configuration of a primary MEC and a secondary MEC according to at least one embodiment disclosed herein, 
         FIGS. 5A-5F  illustrate one configuration of a fault tolerant combined primary and secondary power distribution network of primary MECs, secondary MECS, and a standby MEC according to at least one embodiment disclosed herein, 
         FIG. 6  illustrates one configuration of a secondary power busing network in a forward section of the aircraft according to at least one embodiment disclosed herein, 
         FIG. 7  illustrates one configuration of a MEC for servicing equipment loads and having a computing and network interface module for distributed computing functions and gateway routing of bi-directional data between MECs according to at least one embodiment disclosed herein, 
         FIG. 8  illustrates one configuration of a data network structure with communication bus interfaces between spatially distributed MECs separated by section breaks according to at least one embodiment disclosed herein, 
         FIG. 9  illustrates one configuration of the computing and interface module for distributed computing functions and gateway routing of bi-directional data according to at least one embodiment disclosed herein, 
         FIGS. 10A-10D  illustrate various configurations for high voltage primary power busing structures of primary MECs relative a particular power input source and a plurality of different power outputs according to at least one embodiment disclosed herein, 
         FIG. 11  illustrates a common structure and layout of a primary power switching network device having a common power input source and a plurality of common power outputs for use with the primary MECs according to at least one embodiment disclosed herein, 
         FIGS. 12A-C  illustrate one configuration of a set of primary power switching network devices for use with a primary MEC receiving three-phase power from a generator according to at least one embodiment disclosed herein, 
         FIG. 13  illustrates an exploded perspective view of a multi-layered integrated truss system of a MEC according to at least one embodiment disclosed herein, 
         FIG. 14  illustrates one configuration of a primary MEC with multiple power and communication transfer layers according to at least one embodiment disclosed herein, 
         FIG. 15  generally illustrates one configuration of three-phase primary power routed from main generators to multiple transformer rectifier units (TRUs) and autotransformer units (ATUs) resulting in zero direct current (DC) offset voltage according to at least one embodiment disclosed herein, 
         FIG. 16  illustrates one configuration of the distribution of either alternating current (AC) or DC power from the TRUs and ATUs to equipment loads utilizing twisted and shielded electrical conductor pairs according to at least one embodiment disclosed herein, 
         FIG. 17  illustrates one configuration of an integrated truss system of a MEC within the floor of an aircraft according to at least one embodiment disclosed herein, 
         FIG. 18  illustrates one configuration with integrated protection chipsets (IPCs) where power feeders interconnect with the main generators, the auxiliary power unit, and the MECs for performing local protective functions according to one embodiment disclosed herein, 
         FIG. 19  illustrates one configuration of a MEC having IPCs installed on contactors of a MEC according to at least one embodiment herein, and 
         FIG. 20  illustrates a one configuration of a routine for providing distributed protection and control architecture for electrical systems of a vehicle according to one embodiment disclosed herein. 
     
    
    
     DETAILED DESCRIPTION 
     The following detailed description is directed to vehicles having modular equipment centers to increase vehicle system redundancies while also distributing the modular equipment centers (MECs) throughout the vehicle in such a way that minimizes wire weight and the number of required wire connections to reduce overall vehicle weight and production time. The present invention is susceptible of embodiment in many different forms. There is no intent to limit the principles of the present invention to the particular disclosed embodiments. References hereinafter made to certain directions, such as, for example, “front”, “rear”, “left” and “right”, are made as viewed from the rear of the vehicle looking forward. In the following detailed description, references are made to the accompanying drawings that form a part hereof and in which are shown by way of illustration specific embodiments or examples. Referring now to the drawings, in which like numerals represent like elements throughout the several figures, aspects of the present disclosure will be presented. 
     Aspects of this disclosure may be used in many types of vehicles such as, for example, aircraft, spacecraft, satellites, watercraft, submarines, and passenger, agricultural or construction vehicles. Aspects of this disclosure may also be used in different constructions of vehicles. While the immediate benefit is towards vehicles that have non-conducting frames, chassis or skin, the disclosure features may be suitable and beneficial of vehicles constructed of conductive materials. For the sake of simplicity in explaining aspects of the present disclosure, this specification will proceed utilizing a composite aircraft  10  as the primary example. However, as will be seen, many of aspects of the present disclosure are not limited to the composite aircraft  10 . 
     As well understood by those skilled in the art, the exemplary aircraft  10  depicted in  FIG. 1  includes a fuselage made substantially of composite materials, or composites. The outer composite skin on the fuselage of the aircraft  10  conforms to the curvature of fuselage frames. The fuselage includes a forward section  12 , a middle section  14 , and an aft section  16 . Section breaks  18 ,  20 ,  22  are defined between adjacent aircraft sections. The composite aircraft  10  may have any number of engines. As shown in  FIG. 1 , left engine  30  is supported on the left wing and right engine  32  is supported on the right wing. Each of the engines  30 ,  32  has a rotor which defines a rotor burst zone  38  ( FIG. 5A ) in which damage to the fuselage and aircraft systems between the engines  30 ,  32  may occur as a result of an event or an operational inconsistency with one of the engines  30 ,  32 . 
     The composite aircraft  10  may have any number of sections and the position of the aircraft sections or systems within the composite aircraft  10  may sometimes be described as being forward or aft of the rotor burst zone  38 . Floor beams extend between fuselage frames to define a passenger compartment above the floor beams and a cargo area for holding cargo below the floor beams. Stanchions extending between the fuselage frames and the floor provide a fulcrum to assist in stiffening the floor of the composite aircraft  10 . The passenger area is pressurized and all or part of the cargo area may be pressurized. Ducts may be positioned through the crown run of the composite aircraft  10  above the passenger compartment or below the floor in the cargo area such as between the fuselage frame and the stanchions. 
     On each of the engines  30 ,  32  are one or more main primary power sources such as high voltage AC left power generators  34   a ,  34   b  and high voltage AC right power generators  36   a ,  36   b  (hereinafter may be referred to collectively and/or generically as “left generators  34 ”, “right generators  36 ” or “generators  34 ,  36 ”). Primary power feeders  40   a  and  40   b  extend from the left generators  34   a ,  34   b  and primary power feeders  42   a  and  42   b  extend from the right generator  36   a ,  36   b . As shown in  FIG. 1 , primary power is distributed throughout the composite aircraft  10  via the primary power feeders  40   a ,  40   b ,  42   a ,  42   b  (hereinafter may be referred to collectively and/or generically as “power feeders  40 ,  42 ”). The composite aircraft  10  may also have one or more high voltage AC auxiliary power unit generators  54  for redundancy in the event one or more of the generators  34 ,  36  fail, as well as to provide power when the engines  30 ,  32  are not running. When the composite aircraft  10  is parked and the engines are not running, power may be provided to the aircraft by one or more power sources such as high voltage AC external power unit  56 . 
     For purposes of this disclosure, low voltage and high voltage are those voltages typically referred to as either low or high voltage within the aircraft industry and as may be described in DO-160, Environmental Conditions and Test Procedures for Airborne Equipment, a standard for environmental test of avionics hardware published by RTCA, Incorporated. Throughout this disclosure, 230 VAC is referred to as high voltage but another voltage within a range of voltages, higher or lower than 230 VAC, could also be referred to as high voltage. Also, 28 VDC and 115 VDC are referred to as low voltages but another voltage within a range of voltages, higher or lower than either of 28 VDC and 115 VDC, could also be referred to as low voltage. 
     The composite aircraft  10  in  FIG. 1  does not have dedicated centralized equipment bays for housing power and communications equipment. The equipment is configured into modular power and communication equipment centers, referred to as MECs, that are spatially distributed throughout the composite aircraft  10 . For example, one or more MECs are spatially distributed in each of the forward, middle and aft sections  12 ,  14 ,  16 . Each of the MECs provide localized power conversion and may be either a primary MEC  44 , a secondary MEC  46 , or an auxiliary or standby MEC  48 , as described in greater detail below. Primary MEC  44 , secondary MEC  46  and standby MEC  48  may generally be referred to as “MEC” with one or more applicable reference numbers  44 ,  46 ,  48 . Primary power is distributed from the generators  34 ,  36  via power feeders  40 ,  42  across section breaks  18 ,  20 ,  22  to a primary power input of each of the MECs  44 ,  46 ,  48 . 
     For optimized fault tolerance, the aircraft  10  may include a standby MEC  48  positioned in the rear of the aircraft  10  and at least two MECs  44 ,  46  positioned in each of the forward, middle, and aft sections  12 ,  14 ,  16  of the aircraft  10 . For example, in  FIG. 1  redundancy may be achieved by having multiple MECs  44 ,  46 ,  48  in each aircraft section without having to cross section breaks  18 ,  20 ,  22 . Preferably, each section  12 ,  14 ,  16  includes a primary MEC  44  and a corresponding secondary MEC  46  thereby defining a two by three configuration of MECs  44 ,  46  plus a standby MEC  48 . If there are four separate aircraft sections then there is a two by four configuration of MECs  44 ,  46 . Preferably, the MECS  44 ,  46 ,  48  are alternately spaced on the left and right sides relative to one another along the length of the aircraft  10 . It should be understood that the present disclosure is not limited to any particular number or configuration of MECs  44 ,  46 ,  48 . 
     Equipment loads  50  may be various electrical loads in an aircraft including, but not limited to, displays, fans, environmental units, and the like. Sometimes an equipment load  50  may be in the form of a line replaceable unit (LRU)  52  ( FIG. 4 ). The equipment loads  50  within each of the aircraft sections  12 ,  14 ,  16  are grouped into one or more zones of power and communication. Each zone of equipment loads  50  across multiple systems may be associated with and serviced by the nearest MEC  44 ,  46 . Preferably, each zone of equipment loads  50  is located within a single section and associated with at least one MEC location in the same zone. Preferably, the connecting wires or lines do not cross section breaks  18 ,  20 ,  22 . 
     Generally, any equipment load  50  on the aircraft  10  requires both electrical power and communication data. Data is needed to tell the equipment load  50  what to do, or provide feedback about its current status, while electrical power is needed so the equipment load  50  can perform its intended function. If power and data are provided to an equipment load  50  from different equipment centers and if one of either the power or data is lost then the equipment load  50  then has an indeterminable state. To avoid indeterminate states each MEC  44 ,  46 ,  48  independently provides both the electrical power and communication data for servicing each of the localized equipment loads  50  within an associated zone. The electrical power and data communication to an equipment load  50  may be synced or grouped together in that both the power and the data communication provided to the equipment load  50  originate from a single source such as the nearest MEC  44 ,  46 ,  48 . Synced electrical power and communication data is sometimes referred to as a power channel. Each of the equipment loads  50  within a zone may receive power from a particular MEC  44 ,  46  and therefore the network communication switches providing data to those same equipment loads  50  are powered by that same MEC  44 ,  46 . 
     The MECs  44 ,  46 ,  48  are configured to distribute power received from the main power sources. The MECs  44 ,  46 ,  48  may independently convert the primary power into secondary power. Secondary power may be distributed from the MECs  44 ,  46 ,  48  to then independently service each of the equipment loads  50  within each zone without a secondary branch power network extending across the section breaks  18 ,  20 ,  22 . In such case, control and conversion of the primary power may be distributed to each of the primary MECs  44  of each section of the aircraft  10  such that only primary power is distributed across the section breaks  18 ,  20 ,  22  amongst the primary MECs  44 . In a preferred configuration, only high voltage power feeders and the data backbone cross production breaks. 
     Distributing only primary power across section breaks  18 ,  20 ,  22  reduces the amount of wire required for distributing secondary power across multiple sections of the aircraft  10 . This is because the distributed MEC architecture creates a separate secondary power distribution network within each section that allows for shorter runs of secondary wiring. Doing so reduces the overall weight of the wire utilized throughout the aircraft as well as the number of secondary connections required when joining adjacent fuselage sections. Also, because of the shorter secondary power runs, the total loop area of the power feeder run is reduced as compared to an implementation within a current return network. Moreover, aircraft production processes are improved because the secondary power network of wires extending across section breaks are limited or eliminated. The reduction of secondary power wires extending across section break are more readily tested and build quality verified earlier due to reduced reliance on other sections before final assembly of the aircraft  10 . 
     As shown in  FIG. 1 , primary power feeder  40   a  extends from generator  34   b  on the left engine  30  into the middle section  14  to a MEC  44  shown on the left side of the mid section  14 , across section break  20  to another MEC  44  shown on the left side of forward section  12 , and then to another MEC  44  shown on the left side in front of forward section  12 . Primary power feeder  40   b  extends from generator  34   a  on the left engine  30  into the middle section  14  to a MEC  44  on the left, across section break  22  to a left aft MEC  44 , and then to a left aft MEC  48 . Power feeder  42   a  extends from generator  36   a  on the right engine  32  into the middle section  14 , across section break  20  to a MEC  44  on the right in forward section  12 , and then to another MEC  44  on the right in front of the forward section  12 . Primary power feeder  42   b  extends from generator  36   b  on the right engine  32  into the middle section  14  to middle right MEC  44 , across section break  22  to right aft MEC  44 , and then to right aft MEC  44 . Alternatively, the power feeders  40   a ,  40   b  could instead provide primary power to the MECs  44  on the right side of one or more sections of the aircraft  10 . In such case, the power feeders  42   a ,  42   b  would provide primary power to the MECs  44  on the left side of one or more sections the aircraft  10 . 
     Also, one of the generators  34   a ,  34   b  on the left engine  30  could provide primary power to one side of the aircraft forward of a rotor burst zone  38  and the other of generators  34   a ,  34   b  on the left engine  30  could provide primary power to the other side of the aircraft  10  aft of the rotor burst zone  38 . In such case, one of the generators  36   a ,  36   b  on the right engine  32  could provide primary power forward of the rotor burst zone  38  to the opposite side that is powered by one of the left generators  34   a ,  36   b . The other of generators  36   a ,  36   b  on the right engine  32  could provide primary power aft of the rotor burst zone  38  to the opposite side powered by the other one of the left generators  34   a ,  36   b.    
       FIG. 2  illustrates splitting two generators per engine relative the rotor burst zone  38  of the aircraft  10  which increases the availability of primary power in the event of an operational issue with an engine  30 ,  32 . If one of the engines  30 ,  32  is lost, or a generator  34 ,  36  within one of the engines  30 ,  32  fail, the two remaining generators  34   a ,  34   b ,  36   a ,  36   b  on the remaining engine  30 ,  32  distribute both forward and aft primary power to the aircraft  10 . Generator  34   a  of the left engine  30  and generator  36   a  of the right engine  32  power a pair of primary power switching buses  96   a  forward of the rotor burst zone  38  that are connected to one another by a forward tie bus  76 . Generator  34   b  of the left engine  30  and generator  36   b  of the right engine  32  power another pair of primary power switching buses  96   a  aft of the rotor burst zone  38  that are connected by an aft tie bus  78 . A mid tie bus  80  connects at least one of the forward primary power switching buses  96   a  with at least one of the aft primary power switching buses  96   a . Therefore, when an engine  30 ,  32  experiences an operational inconsistency, the aircraft  10  continues to have power and control on one side along the entire length of the aircraft  10  due to the distribution of power from the remaining engine  30 ,  32  in a forward and aft manner. The power and control is distributed from a single engine  30 ,  32  both forward and aft of the rotor burst zone  38  without increasing the amount of wiring.  FIG. 2  also illustrates the primary power switching buses  96   a  distributing power to the secondary MECs  46  for power conversion and distribution to equipment loads  50  as explained in greater detail below. A standby MEC  48  may be coupled to the secondary MECs  46  to provide backup power when the primary main AC power sources are not available to the primary power switching buses  96   a  as explained in greater detail below. 
     Unserviced equipment loads  50  in one or more zones occurs primarily for two reasons. Either all of the generators  34 ,  36  failed and therefore primary power is no longer available to any of the MECS  44 ,  46  or one or more of the buses  96  are physically damaged due to an event such as a rotor or tire burst. Rerouting of high voltage power from either of the four generators  34 ,  36  or auxiliary power unit generator  54  based on failure of one or more main primary power sources occurs at the primary bus level via the tie buses  76 ,  78 ,  80  through opening and closing of combinations of switches as shown by the primary power busing network system  90  depicted in  FIG. 3 . In one or more embodiments, one or more standalone solid state switches, for example contactors, are included on the primary power switch network system  90 . The solid state switches each have a self-contained control function configured to provide one or more of localized protections, voltage sensing, and current sensing, independent of the availability of other power system components. The standalone solid state switch can function without the need for data from other power system components. Opening and closing of the solid state switch interrupts and routes primary power across one or more of the primary power switching buses to one or more of the MECs  44 ,  46 ,  48 . Beginning with  FIG. 3 , specific contactors are depicted as either primarily closed or primarily open. The symbol for an open contactor is two parallel lines. The symbol for a normally closed contactor is the same with the exception that a diagonal line is drawn through the parallel lines. The standalone solid state switch may also include pulse width modulation to limit current flow through the standalone solid state switch. Rerouting of secondary power and low voltage DC between the MECs  44 ,  46 ,  48  based on failure of high voltage buses and conversion occur by the opening and closing of combinations of switches as shown by the primary power busing network  90  as depicted in  FIG. 3 . 
     Each MEC  44 ,  46 ,  48  has both primary and secondary power and is capable of independently performing closed loop processing and local control of sensors without being dependent on a central computer system. The distributed power system control architecture permits sharing of the overall vehicle power distribution status among the MECs  44 ,  46 ,  48  but each MEC  44 ,  46 ,  48  is only responsible for servicing equipment loads  50  in proximity of each MEC, with the exception of MEC  48  which also distributes standby power to all other MECs  44 ,  46 . Each MEC  44 ,  46 ,  48  manages data associated with the zone of the nearest equipment loads  50  such that each MEC  44 ,  46 ,  48  independently performs operations within its own zone of equipment loads  50 . 
     Each MEC  44 ,  46 ,  48  also preferably has solid state switching for bus power control and also provides circuit protections. In  FIG. 3  power from primary power feeders  40 ,  42  connected to the generators  34 ,  36  energizes primary power switching buses  96   a . Each primary power switching bus  96   a  branches off to a primary power switching bus  96   b  within MEC  44  and a primary power switching bus  96   c  within MEC  46 . Each primary power switching bus  96   a  connected with distribution feed  98  to a primary power switching bus  96   b  corresponds with a single primary MEC  44  as shown in  FIG. 4  and as described in greater detail below. 
     Referring to  FIG. 4 , a portion of each primary MEC  44  with a primary power switching bus  96   a  is a high power portion  120  and another portion of the primary MEC  44  with the primary power switching bus  96   b  is a low power portion  122  of the primary MEC  44 . The high power portion  120  of a primary MEC  44  is configured to receive primary power from any high power main source available to the aircraft  10  and is sometimes referred to as a primary power switching network device  302  ( FIG. 12A-12C ). The network of high power portions  120  of the primary MECs  44  within the aircraft  10  define a high voltage primary power switching network. 
     The low power portion  122  is preferably configured to handle a fraction of the power from onboard power sources but still be able to handle the same voltages as the high power portions  120 . The primary power switching buses  96   c  correspond with secondary MECs  46  shown in  FIG. 4 .  FIG. 4  best illustrates the similarity between a secondary MEC  46  and the low power portion  122  of a primary MEC  44 . Primary MECs  44  include the primary level power network busing structure of primary power switching buses  96   a  to reroute primary sources across the aircraft  10  that the secondary MECs  46  do not have. During normal as well as abnormal operations, the primary and secondary MECs  44 ,  46  both have primary and standby power. Secondary MECs  46  service the nearest equipment loads  50  just like a primary MEC  44 . 
     Referring back to  FIG. 3 , distribution feeds  98  extend between primary power switching buses  96   a  and  96   b  of each primary MEC  44  and distribution feeds  100  extend between each bus  96   b  of the primary MEC  44  and the primary power switching bus  96   c  of a secondary MEC  46  that directly receives power from the same source. Also, a crosstie  102  extends between the bus primary power switching  96   b  of the primary MEC  44  associated with left generator  34   a  and the primary power switching bus  96   b  of the primary MEC  44  associated with the right generator  36   a . A crosstie  104  extends between the primary power switching bus  96   c  of the secondary MEC  46  associated with left generator  34   a  and the primary power switching bus  96   c  of the secondary MEC  48  associated with the right generator  36   a . A crosstie  106  extends between the primary power switching bus  96   b  of the primary MEC  44  associated with left generator  34   b  and the primary power switching bus  96   b  of the primary MEC  44  associated with the right generator  36   b . A crosstie  108  extends between the primary power switching bus  96   b  of the secondary MEC  46  associated with generator  34   b  and the primary power switching bus  96   b  of the secondary MEC  46  associated with the right generator  36   b . Auxiliary power unit generator  54  is connected to the crossties  102 ,  106 , respectively. 
       FIG. 5A  illustrates one configuration of the fault tolerant combined primary and secondary power distribution networks of primary, secondary and standby MECS  44 ,  46 ,  48  within the aircraft  10 . For purposes of illustrating greater detail,  FIGS. 5B-5E  illustrate close-up partial views of four separate portions which can be positioned next to each other to assemble the complete system depicted in  FIG. 5A . Two dot-dash lines on each of  FIGS. 5B-5E  denote the broken edges of each partial view.  FIG. 5B  illustrates the top left portion of  FIG. 5A .  FIG. 5C  illustrates the top right portion of  FIG. 5A .  FIG. 5D  illustrates the bottom left portion of  FIG. 5A  and  FIG. 5E  illustrates the bottom right portion of  FIG. 5A . Also,  FIG. 5F  illustrates one configuration of the standby MEC  48  of the system of  FIG. 5A . The contactors shown in  FIG. 3  are also shown symbolically in  FIGS. 5A-5F , but without reference numbers to simply  5 A- 5 F, and may also be shown in other drawings without any reference numbers or having different reference numbers. 
     In  FIG. 5A  the primary and secondary MECS  44 ,  46  are arranged in such a way that there are a total of four in the forward sections of the aircraft  10  and another four in the aft sections of the aircraft  10 . Preferably, there is a primary MEC  44  and a secondary MEC  46  in each of a pair of forward sections and a primary MEC  44  and a secondary MEC  46  in each of a pair of aft sections.  FIG. 5A  also shows a standby MEC  48  in an aft section of the aircraft  10 . The non-time limited power source for the standby MEC  48  can be a RAM air turbine (RAT)  128  or other suitable independent time limited standby power source such as a battery or fuel cell. In the event of an operational inconsistency with all the generators  34 ,  36 , the RAT  128  is deployed to provide standby power to standby MEC  48  as well as to one or more of the MECs  44 ,  46  in the event that all of the generators  34   a ,  34   b ,  36   a ,  36   b  have an operational inconsistency. The battery  598  provides temporary operational power to standby MEC  48  as well as to one or more of the MECs  44 ,  46  while the non-time limited RAT  128  is being deployed. 
     If one of the generators  34   a ,  34   b ,  36   a ,  36   b  fails, power is not being received at the primary power switching bus  96   a  of a primary MEC  46 . Therefore, the equipment loads  50  off of the lower power portion  122  of the primary power switching bus  96   b  of the unpowered primary MEC  44  are unserviced and the equipment loads  50  off of the primary power switching bus  96   c  of an unpowered adjacent secondary MEC  46  are unserviced. Power is then rerouted at the primary level from one of the other remaining operational sources by opening and closing of combinations of contactors to energize primary power switching bus  96   a  of unpowered primary MEC  44  to power its equipment loads  50  and to energize primary power switching bus  96   c  of any unpowered adjacent secondary MEC  46  to power its equipment loads  50 . 
     Alternatively, if a MEC  44 ,  46 ,  48  experiences a physical failure and as result its equipment loads  50  are unpowered, then power may be rerouted to power the equipment loads  50  of the unpowered MEC  44 ,  46 ,  48  by another powered MEC  44 ,  46 ,  48 . Depending on the amount of power available to be rerouted, all or only a portion of the equipment loads  50 , such as only the critical loads, may be repowered. Also, if all power sources are lost and the MECs  44 ,  46 ,  48  are unpowered, then the standby MEC  48  with the fuel cell or RAT  128  can power the critical equipment loads  50  of the other MECs  44 ,  46 . Critical loads are those equipment loads  50  that the aircraft  10  must have powered to maintain continued safe flight and landing. Essential loads are those equipment loads  50  that are desirable to have such as radios and other communications equipment but operation is not required to fly the aircraft  10 . Non-essential loads are the lowest priority equipment loads  50  such as passenger comfort loads including food preparation devices, decorative lighting and cabin entertainment systems. 
     By way of example, the auxiliary power unit generator  54  could service the equipment loads  50  lost due to the failure of one of the main generators  34 ,  36 . If generator  34   b  fails then, through a combination of contactors in forward tie bus  76 , aft tie bus  78 , mid tie bus  80 , primary power is provided directly from the remaining main generators  34 ,  36 . Alternatively, primary power may be provided from the auxiliary power unit generator  54  through another operational MEC  44 ,  46  across one or more of the crossties  102 ,  104 ,  106 ,  108 , to the primary power switching bus  96   a  of an unpowered primary MEC  44  or to the primary power switching bus  96   c  of an unpowered secondary MEC  46 . 
     In the event one or more of the MECs  44 ,  46  has a physical operational inconsistency, all or part of the plurality equipment loads  50  within the zone associated with each operationally inconsistent MEC  44 ,  46  can be associated with one or more other MECs  44 ,  46  that are nearest in proximity. For example, if a primary MEC  44  physically fails, the equipment loads  50  once serviced by that failed MEC  44  may be serviced by another MEC  44 ,  46  or a combination of MECs  44 ,  46 . MECs  44 ,  46  can determine the types of equipment loads  50  once serviced by the failed MEC  44  and then determine whether one or more of the combination of MECs  44 ,  46  should service those unpowered equipment loads  50 . If it is determined that a secondary MEC  46  in closest proximity to the failed primary MEC  44  is to service the additional equipment loads  50  then the zone originally associated with that secondary MEC  46  is expanded to encompasses the zone formerly serviced by failed primary MEC  44 . 
     Alternatively, the additional equipment loads  50  may be divided between a secondary MEC  46  and another primary MEC  46  in proximity to the failed primary MEC  46 . In such case, the zone of equipment loads  50  associated with the nearest operational primary MEC  44  is expanded to include a portion of the zone formerly serviced by failed primary MEC  44  and the zone of equipment loads  50  associated with the nearest operational secondary MEC  46  is expanded to include the remaining portion of the zone formerly serviced by failed primary MEC  44 . In either case, one or more other MECs  44 ,  46  in proximity of a failed MEC  44 ,  46  are sourced to independently provide the services to the equipment loads  50  previously serviced by the failed MEC  44 ,  46 . 
     Each secondary MEC  46  and each low power portion  122  of each primary MEC  44  includes contactors coupled to conversion equipment. The conversion equipment includes a transformer rectifier unit (TRU)  134 , which rectifies the 230 VAC and converts it to the main DC output such as 28 VDC for bus  136 , and an autotransformer or autostep down transformer unit (ATU)  138  to convert 230 VAC to 115 VAC for a low power AC output bus  140 . Each secondary MEC  44  and low power portion  122  of a primary MEC  44  further includes a second TRU  142 , not just for redundancy, but to provide power only to the critical loads absolutely necessary for continued safe flight and landing. Limiting the second TRU  142  to only critical loads ensures that the standby power sources are not overloaded. 
       FIG. 6  illustrates the configuration of a secondary power busing configuration, in forward section  12  for example, where the primary power switching buses  96   b  in the low power portions  122  of the primary MECs  44  and the primary power switching buses  96   c  of the secondary MECs  46  are tied together. As described above, whether all or only a portion of the unpowered equipment loads  50  of a damaged MEC  44 ,  46  are serviced by another MEC  44 ,  46  depends on available power. In the event one of the TRUs  134  in one of the MECs  44 ,  46  within an aircraft section fails, the most critical of equipment loads  50  from the operationally inconsistent TRU  134  may be serviced by another MEC  44 ,  46  in that same aircraft section providing secondary power across the various contactors and backup buses  148 . 
     Preferably MECs  44 ,  46  in the aft section  16 , have secondary power tie-ins from the auxiliary power unit generator  54  due to their proximity to one another which minimizes the power feeder wire weight. Also, the MECs  44 ,  46  in the forward section  12  of the aircraft  10  tie in at lower voltage levels such as 115 VAC from the external power ground service equipment such as external power unit  56  as shown in  FIGS. 2 and 6 . However, the 115 VAC from the ground to the low power AC output buses  140  in the MECs  48  in the forward section  12  could be converted to higher voltages such as 230 VAC by bi-directional ATUs  138  which then may be distributed to the other MECs  44 ,  46  in other sections of the aircraft  10 . Also, a second TRU  142 , typically used for more critical loads as explained above, allows battery power from battery bus  294  via backup bus  148  to power those critical loads that were lost. 
     As shown in  FIG. 7 , a computing (hardware and software) and network interface (CNI) module  162  located inside each MEC  44 ,  46 ,  48  provides distribute computing functions and gateway routing of bi-directional data. Each CNI module  162  contains two failsafe computing systems that become a fault tolerant computing system. Each fail safe computing system is redundant to the other. This fault tolerant computing system responds gracefully to unexpected hardware and/or software failures to allow no loss of service to system functions within the aircraft  10 . The CNI module  162  transmits/receives data to/from internal MEC computing functions and external MEC computing functions via an internal system communication bus (such as FlexRay, Controller Area Network (CAN), ARINC  664 , TTP or other bus technology). Other MECS  44 ,  46 ,  48  on the aircraft  10  will communicate with the CNI module  162  via a data networking specification, such as the ARINC  664 , across external data communication channel A and external data communication channel B as shown in  FIG. 7  having reference numbers  188  and  190 , respectively. 
     The CNI module  162  is a distributed computing element that hosts specific software applications used within that localized zone of the aircraft  10 . Some examples of the system applications that can be hosted on the CNI module  162  are the AC and DC power systems, cargo door system, passenger entry door system, landing gear system, and passenger cabin system. The computing functions that communicate to the CNI module  162  are TRUs  134 , TRUs  142 , ATUs  138 , solid state switches of a breakers module  166 , a generator control unit GCU  168  associated with one of the generator  34 ,  36 , solid state power distribution modules  170 , and remote data concentrators. The CNI module  162  communicates internally within the MEC  44 ,  46 ,  48  across internal data channel A  202  and internal data channel B  204  to the TRUs  134 ,  142 , the ATUs  138 , the breaker modules  166 , the GCU  168 , and the power distribution modules  170  as described in greater detail below. 
     The CNI module  162  will transmit and receive data to/from these computing functions. The CNI module  162  will also transmit and receive the status and health from other MECs  44 ,  46 ,  48  and aircraft computing systems. Each CNI module  162  manages the workload of an individual MEC  44 ,  46 ,  48  with knowledge of what is going on in other MECs  44 ,  46 ,  48 . Once the information has been received by the CNI module  162  of a MEC  44 ,  46 ,  48 , its computing function will determine which system needs the data, interpret the health of the data, respond to any power system anomalies, supply time-critical information to computing functions that need it, perform system level logic algorithms, report airplane level system faults, and control the distribution of AC and DC power for that zone. 
       FIG. 8  illustrates the data network structure with communication bus interfaces between the spatially distributed MECs  44 ,  46 ,  48  separated by section breaks  18 ,  20 ,  22 . This configuration allows each individual MEC  44 ,  46 ,  48  to communicate with other MECs  44 ,  46 ,  48  as well as provide the redundancy required to ensure continued communication across failures. Section break  20  defines forward and aft sections of the aircraft. The number of network communication switches needed is determined by the number of MECs  44 ,  46 ,  48  and the desired fault tolerance.  FIG. 8  illustrates nine MECs  44 ,  46 ,  48  with three pairs of network switches  182   a - b ,  184   a - b ,  186   a - b  (hereinafter may be referred to collectively and/or generically as “network switches  182 ,  184 ,  186 ”). Each network switch  182 ,  184 ,  186  may be a multilayer network switch such as a layer-3 network switch that can receive secondary electrical power from the CNI modules  162  of each of the interfacing MECs  44 ,  46 ,  48 . If there were more MECs  44 ,  46 ,  48  then more network switches would be required to achieve the same level of fault tolerance. 
     Each MEC  44 ,  46 ,  48  has A and B communication channels. Channels A and B of each primary MEC  44  connects to two corresponding A or B switches on either another primary MEC  44  or on a standby MEC  48 . Each primary MEC  44  includes one switch  182 ,  184 ,  186  on either channel A or channel B, while standby MEC  48  in an aft section of the aircraft includes both switches of a pair of switches  182 ,  184 ,  186  on both A and B channels. Switches  182   a ,  184   a ,  186   a  correspond with channel A and switches  182   b ,  184   b ,  186   b  correspond with channel B. External communication data lines  192  indicate switch to switch data lines. 
     Generally, a network switch on each primary MEC  44  on one side of the section break  20  is connected to two other network switches of other primary or standby MECs  44 ,  48  where at least one of those MECs  44 ,  48  is on the other side of the section break  20  and one is on the opposite side of the aircraft  10 . For example, network switch  182   a  of the forward right primary MEC  44  that is forward of the section break  20  is connected on the other side of the section break  20  to both network switch  184   a  on the aft left primary MEC  44  and to network switch  186   a  on the standby MEC  48 . Network switch  182   b  on the forward left primary MEC  44  that is forward of the section break  20  is connected on the other side of the section break  20  to both network switch  184   b  on the aft right primary MEC  44  and to network switch  186   b  on the standby MEC  48 . Network switch  186   b  on the standby MEC  48  is also connected to network switch  184   b  on the opposite side of the aircraft  10 . Network switch  184   a  is also connected to network switch  186   a  of the standby MEC  48 . 
     Each of the secondary MECs  46  also has two data channels with two other primary or standby MECs  44 ,  48 . External communication data lines  196  indicate data connections of a network switch of a primary MEC  44  directly to a secondary MEC  44 . One of the channels of each secondary MEC  48  is connected to a network switch on the same channel of a primary MEC  48  on the other side of the section break  20  and the other channel is connected to another secondary MEC  46 . Therefore,  FIG. 8  shows eight data bus connections crossing section break  20  and four data bus connections crossing each of the section breaks  18 ,  22 . This configuration minimizes the amount of communication wiring across section breaks as well as the overall weight of wiring in the aircraft. Separation is maintained between each data bus by utilizing the space in the crown and the floor of the aircraft  10 . 
     If any two MECS  44 ,  46 ,  48  are powered then the communication network will be active and data will be present so that those two MECS  44 ,  46 ,  48  can fully communicate with each other. This communication network is a fault tolerant network in that any one connection between a pair of MECs may be lost without reducing any MEC  44 ,  46 ,  48  functionality. Moreover, loss of any two communication connections at the same time between the MECs  44 ,  46 ,  48  at most results in the loss of data communication with only one of the MECs  44 ,  46 ,  48 . Healthy CNI modules  162  can optimally respond to changing configurations of the power system in a coordinated fashion by utilizing local environment information and communications from other healthy CNI modules  162 . 
     For example, loss of the network switch  182   a  on channel A of the forward right primary MEC  44  does not result in complete loss of communications to and from the forward right primary MEC  44  because communications to and from forward right primary MEC  44  may continue through channel B. Any other MECs  44 ,  46 ,  48  which had communicated via channel A with the forward right primary MEC  44  can directly communicate through channel B or via other MECs  44 ,  46 ,  48  that are connected to forward right primary MEC  44  via channel B. Also, if network switch  182   a  on channel A of the forward right primary MEC  44  was lost in addition to the channel B connection to the forward right secondary MEC  44 , communications to and from the forward right primary MEC  44  would continue via channel B but then communications would be lost only with the forward right secondary MEC  44  because both channels A and B were lost. 
     One aspect of the present disclosure is distributed power control architecture. Power control is distributed to each MEC  44 ,  46 ,  48  as well as power itself. Based on the local data each individual MEC  44 ,  46 ,  48  collects, each MEC  44 ,  46 ,  48  performs its own power control of its associated zone to configure its own equipment loads  50  without having to rely on any other MECs  44 ,  46 ,  48 . Only the data that is really necessary, such as the need to reroute power, is sent to the CNI modules  162  of other MECs  44 ,  46 ,  48 . 
     Normal power up of an aircraft  10  on the ground is preferably a sequential power up of the MECs  44 ,  46 ,  48 . Normal power up is done via the battery  598  which powers all the standby buses  160  in MECs  44 ,  46  via the static inverter  290  and the backup bus  148 . Should the battery  598  not be available, a limited amount of exterior power from the external power unit  56  is sent to power up the standby MEC  48 . Once the standby MEC  48  is powered up, power is then distributed from the standby MEC  48  to the each of the other primary and secondary MECs  44 ,  46  to power up their CNI modules  162  and configure contactors within each MEC  44 ,  46  as appropriate with the power sources that are available. On the other hand, a sequential power up is not utilized if a MEC  44 ,  46  becomes unpowered during normal flight operations. If the CNI module  162  in one of the MECs  44 ,  46  has no primary power, the low power interconnection between two MECs  44 ,  46 , such as a primary MEC  44  and a secondary MEC  44  with a distribution feed  100 , provides a means to still power the unpowered MEC  44 ,  46  as explained above. 
     The CNI module  162  reads input/output communications from other systems or LRUs as well as configuration data from other MECs  44 ,  46 ,  48 . Broadcasting each MEC&#39;s  44 ,  46 ,  48  configuration data allows each of the other MECs  44 ,  46 ,  48  to determine what is going on elsewhere in the aircraft  10 . The CNI module  162  then uses this data to configure breakers and contactors within its MEC  44 ,  46 ,  48  and then writes configuration data onto channel A or B about the equipment loads  50  within its zone to broadcast to the other MECs  44 ,  46 ,  48  so that other MECs  44 ,  46 ,  48  can do the same. Each CNI module  162  checks the validity of the communications input/output and environmental data it receives and, if necessary, refines it to determine its own environment data and states of its breakers. Once the CNI module  162  figures out how it wants to command its breakers and contactors within its zone, it then sends its configuration data out to the other MECs  44 ,  46 ,  48 . 
     The CNI module  162  of each MEC  44 ,  46 ,  48  only controls the equipment loads  50  within the boundaries assigned to that MEC  44 ,  46 ,  48 . Each CNI module  162  of a particular MEC  44 ,  46 ,  48  does not set the equipment load  50  configuration of other MECs  44 ,  46 ,  48  or how to configure their breakers or contactors. However, all the MECs  44 ,  46 ,  48  still cooperate with one another to provide a coherent and unified power transfer function for the primary and secondary power systems of the aircraft  10 . The CNI modules  162  of MECs  44 ,  46 ,  48  that are functioning properly are able to react to a MEC  44 ,  46 ,  48  that has operational issues and reroute power across power tie buses  76 ,  78 ,  80 , distribution feeds  98 ,  100  and crossties  102 ,  104 ,  106 ,  108  even in conjunction with additional failures. The computing and networking architecture is both fail safe and fault tolerant. If a CNI module  162  has an operational issue, all of its connected load will enter a predefined default “fail safe” state. Adjacent CNI modules  162  do not have the capacity or authority to control other equipment loads outside of their zone. 
     The CNI module  162  shown in  FIG. 9  includes one network switch  182 ,  184 ,  186  on one side of the CNI module  162  corresponding with channel A and another network switch  182 ,  184 ,  186  on the other side corresponding with the channel B. Both network switches  182 ,  184 ,  186  have one or more ports  206  for making external data communication connections. Each side of the CNI module  162  also has one or more ports  208  for making internal data communication connections within the MEC  44 ,  46 ,  48  as described in greater detail below. The CNI module  162  includes two multi-core processors  242 ,  244  for running multiple instructions associated with processing channel A and channel B data communications. Each processor  242 ,  244  can process instructions for receiving and transmitting communication data within a MEC  44 ,  46 ,  48  at ports  208  or for receiving and transmitting communication data outside of the MEC  44 ,  46 ,  48  with either network switch  182 ,  184 ,  186  through the ports  206 . One of the processors  242 ,  244  on one side of the CNI module  162  corresponds with one communication channel and the other processor  244  on the other side of the CNI module  162  corresponds with the other communication channel. However, each processor  242 ,  244  has a crossover to the other network switch  182 ,  184 ,  186  for the other communication channel so that each processor  242 ,  244  can read and process both channel A and B communications. 
     Each component or LRU  52  placed on a truss system of a MEC  44 ,  46 ,  48  such as the CNI module  162  may include a barcode reader  248  for optically reading labels. The barcode reader  248  may be a quick response (QR) code reader for reading QR codes. Barcodes (not shown) may be placed in the MEC,  44 ,  46 ,  48  or elsewhere in the aircraft  10  in proximity of the barcode reader  248 . The barcode reader  248  reading barcodes allows the MEC  44 ,  46 ,  48  to input information such as identification, position, time tracking and other configuration information to set software parameters of the CNI module  162  of the MEC  44 ,  46 ,  48 . For example, the bar code reader  248  may read the position of the CNI module  162  so that the MEC  44 ,  46 ,  48  knows which section or which side of the aircraft  10  it is located in. Also, determining the location of the CNI module  162  allows the MEC  44 ,  46 ,  48  to determine the nearest equipment loads  50 . The configuration information may also be transmitted to other MECs  44 ,  46 ,  48 , elsewhere in the aircraft  10 , or a central facility outside of the aircraft  10  such as a maintenance facility. 
     Based on how much power is distributed from the MEC  44 ,  46 ,  48 , the CNI module  162  may require one or more additional power inputs  288 , such as 28 VDC or 115 VAC, and power regulators  238 , from one or more transfer layers of a truss system as explained below. For example, 28 VDC is input to point of use regulator  280  for the barcode reader  248 . Each CNI module  162  also receives one or more DC power inputs  284  from power outputs  286  of the CNI modules  162  of one or more other MECs  44 ,  46 ,  48  to power one or both network switches  182 ,  184 ,  186 . Power inputs  284  and power regulators  246  provide redundancy to prevent a single power failure from resulting in failure of any of the processing or communication channels. 
     If there is a complete loss of power to a MEC  44 ,  46 ,  48  at inputs  288  from a transfer layer of the truss system, then the MEC  44 ,  46 ,  48  with the CNI module  162 , network switches  182 ,  184 ,  186 , the power regulators  246 , and the barcode reader  248 , may still be powered. Because of the one or more DC power inputs  284  routed from redundant power outputs  286  of other CNI modules  162  of other MECs  44 ,  46 ,  48 , the CNI module of the unpowered MEC  44 ,  46 ,  48  never loses power and is able to reroute power from an adjacent MEC and then powers up one or more transfer layers of its own MEC  44 ,  46 ,  48 . The MEC  44 ,  46 ,  48  can then still service some or all of its equipment loads  50  and the CNI module  162  remains fully functional and can communicate with other CNI modules  162  thereby keeping truss system of the MEC  44 ,  46 ,  48  and the communications network active. 
       FIGS. 10A-10D  illustrate configurations of the high voltage primary power switching bus structure of each of the primary MECs  44  shown in  FIGS. 5A-5E . Each may be designated as R1, R2, L1 or L2 based on which generator  34   a ,  34   b ,  36   a ,  36   b  directly powers each of the four primary MECs  44  and whether the four primary MECs  44  are forward or aft and whether they are on the left or on the right side of the aircraft  10 . R1 corresponds with the forward right primary MEC  44  that receives primary power from generator  36   a . R2 corresponds with the aft right primary MEC  44  that receives primary power from generator  36   b . L1 corresponds with the forward left primary MEC  44  that receives primary power from generator  34   a . L2 corresponds with the aft left primary MEC  44  that receives primary power from generator  34   b.    
       FIG. 10A  illustrates the high power portion  120  with primary power switching bus  96   a  and solid state switching devices of the forward right primary MEC  44  (R1) of  FIGS. 5A and 5B .  FIG. 10B  illustrates the high power portion  120  with primary power switching bus  96   a  and solid state switching devices of the aft right primary MEC  44  (R2) of  FIGS. 5A and 5C .  FIG. 10C  illustrates the high power portion  120  with primary power switching bus  96   a  and solid state switching devices of the forward left primary MEC  44  (L1) of  FIGS. 5A and 5D .  FIG. 10D  illustrates the high power portion  120  with primary power switching bus  96   a  and solid state switching devices of the aft left right primary MEC  44  (L2) of  FIGS. 5A and 5E . Together,  FIGS. 10A-10D  depict a common architecture and layout of solid state switching devices that can provide connectivity for each primary MEC  44 . 
       FIG. 10A  best illustrates the primary busing structure, solid state elements, and connections for the forward right primary MEC  44  (R1) which has the least number of solid state elements in comparison to the other primary MECs  44  positioned elsewhere. However, the minimal structure depicted in  FIG. 10A  may be expanded to include the additional solid state elements (shown in phantom) in order to include the required functionality for any of the other primary MECs  44 . The additional solid state elements may or may not be populated in all slots in all installed MECs  44 ,  46 ,  48 . 
     Each of the four configurations of  FIGS. 10A-10D  for the four primary MECs  44  have a primary power connection  210  from one of the main generators  34   a ,  34   b ,  36   a ,  36   b , and a connection  212  to either the forward or aft ties  76 ,  78 . Each configuration also includes an output connection  214  to an associated secondary MEC  46 . Each also includes two high current solid state contactors  216 ,  218  and two low current solid state contactors  220 ,  222 . The two high current contactors  216 ,  218  are connected together at connection  224 . One of the high current contactors  216  is also connected at connection  210  for turning on and off main primary power and the other high current contactor  218  is also connected at the connection  212  for the forward or aft ties  76 ,  78  depending on whether the primary MEC  44  is in a forward or aft section of the aircraft  10 . The low current contactor  220  is connected to a connection  214  for the associated secondary MEC  46 . The other low current contactor  222 , in combination with the distribution feed  98  as described in greater detail below, is for turning on and off power between the high power portion  120  and the low power portion  122  of each primary MEC  44 . 
     The left forward primary MEC  44  (L1) depicted in  FIG. 10C  includes another high current contactor  250  between a connection  252  from the mid tie  80  and the connection  212  for the forward tie bus  76 . The aft left primary MEC  44  (L2) depicted in  FIG. 10D  includes the additional high current contactor  250  that the left forward primary MEC  44  (L1) includes as well as another high current contactor  260  between the connection  252  for the mid tie bus  80  and an input connection  262  for the auxiliary power unit generator  54 . The aft left primary MEC  44  (L2) also includes the same low current contactor  232  to the standby MEC  48  where high voltage AC power is sent across tie  270  and to auto-transformer rectifier unit (ATRU)  272  and to bus  274 . 
     The aft right primary MEC  44  (R2) depicted in  FIG. 10B  also includes a low current contactor  232  for connecting the standby MEC  48  with high voltage AC power across tie  234  to ATRU  236  and to bus  240 . All four configurations also have the option of having additional contactors such as a low current contactor  278  for turning on and off power requiring 230 VAC as shown in each of the  FIGS. 10A-10D . 
     To facilitate manufacturing and inventory availability, the varying architectures shown in each of the  FIGS. 10A-10D  can be rearranged into a single structure having a similar layout of a primary power switching network device (PPSND)  302 , as shown in  FIG. 11 , with optional contactors  232 ,  250 ,  260 ,  278  for various loads based on the primary power switching configuration and where the MEC  44  is within an aircraft  10  as explained above. Each PPSND  302  corresponds with the high power portion  120  of each primary MEC  44  and is configured to share common sources and outputs with options for additional contactors  232 ,  250 ,  260 ,  278  for receiving primary power directly from a standby MEC  48  or for receiving primary power from auxiliary power unit generator  54  which are connected via the forward, aft and mid tie buses  76 ,  78 ,  80  as needed. 
     As shown in  FIG. 5C  the high power primary power switching bus  96   a  of aft right primary MEC  44  is connected to the standby MEC  48  with tie  234 . In  FIG. 5D , the forward left primary MEC  44  is connected by the forward tie bus  76  with contactor  218   a - c  to the forward right primary MEC  44  and connected by the mid tie bus  80  with contactors  250   a - c  to the aft left primary MEC  44 . The aft left primary MEC  44  as shown in  FIG. 10D  has the most contactors as a result of being connected to other primary MECs  44  with the aft and mid tie buses  78 ,  80  as well as the standby MEC  48  with contactors  232   a - c  by the tie  270 . 
     A set of substantially identical PPSNDs  302   a - c  as shown in  FIGS. 12A-12C , can be used with a primary MEC  44  to receive three-phase primary power from one of the generators  34 ,  36 . Although the PPSNDs  302   a - c  shown are labeled for use in combination with the forward right primary MEC  44  (R1), the three PPSNDs  302   a - c  may also be used to receive three-phase power for either of the other primary MECs  44 . Each of the primary power feeders  40 ,  42  preferably is a four conductor power wire connected to each of the primary MECs  44  where three of the conductors carry either of phases A, B or C of the three-phase power. The forth conductor can be a neutral wire connected to a fourth PPSND. 
     Still referring to  FIG. 11  and  FIGS. 12A-12C , phase A power is received at connection  210   a  to power the primary power switching bus  96   a  of the PPSND  302   a  of  FIG. 12A , phase B power is received at connection  210   b  to power the primary power switching bus  96   b  of the PPSND  302   b  of  FIG. 12B , and phase C power is received at connection  210   c  to power the primary power switching bus  96   c  of the PPSND  302   c  of  FIG. 12C . Solid state elements are depicted by squares in each of  FIGS. 12A-12C  and sets of the solid state elements constitute the contactors  216 ,  218 ,  220 ,  222 ,  232 ,  250 ,  260  and  278  as depicted in  FIG. 11 . 
     In  FIGS. 12A-12C , reference numbers ending with “a”, “b” or “c” may refer to components utilizing phase A power, phase B power or phase C power, respectively. However, such reference numbers themselves may also refer collectively and/or generically to the same components without specifically referencing a particular phase of power. Connections  210   a - c  power the high power primary power switching buses  96   a - c , respectively. Primary power comes from each of the primary power switching buses  96   a - c  across contactors  218   a - c  to the forward tie bus  76   a - c  (or the aft tie bus  78   a - c  depending on whether the primary MEC  44  is in a forward or aft section of the aircraft  10 ). Alternatively, power could come from the forward tie bus  76   a - c  to the primary power switching bus  96   a - c  across contactors  218   a - c . Primary power could also be provided to and from the mid tie bus  80   a - c  at connections  252   a - c  and across contactors  250   a - c  in association with the primary power switching buses  96   a - c . Primary power could also be provided from the auxiliary power unit generator  54  with a power tie  130  connected to connection  262   a - c  and across contactors  260   a - c  to the primary power switching buses  96   a - c.    
     Primary power is provided from the primary power switching buses  96   a - c  across contactors  220   a - c  to the output connections  214   a - c  for the secondary MEC  46 . Primary power is also provided from the primary power switching buses  96   a - c  across contactors  222   a - c  to output connections  390   a - c  and across distribution feed  98  to power the low power portion  122  of the primary MEC  44 . Three-phase primary power from the output connections  390   a - c  of the PPSNDs  302   a - c  may be sent through a truss system to other components within the same MEC  44 ,  48  as the PPSNDs  302   a - c . Distribution feed  98  is preferably a four wire conductor with a first wire for phase A power connected to output connection  390   a , a second wire connected to output connection  390   b , and a third wire connected to output connection  390   c.    
     Three-phase high power may be distributed directly from the high power portion  120  of a primary MEC  44  to optional or auxiliary loads by utilizing output connections  340   a - c  or output connections  342   a - c  on the PPSNDs  302   a - c  shown in  FIGS. 12A-12C . Contactor  232  and contactor  278 , shown in  FIG. 11 , correspond with output connections  340   a - c  and output connections  342   a - c  shown in  FIGS. 12A-12C . Power of 230 VAC to and from the auxiliary loads is controlled by the contactors  232   a - c ,  278   a - c  of the PPSNDs  302   a - c . If the PPSNDs  302   a - c  were being utilized in the left aft primary MEC  44  (L2) as shown in  FIGS. 5A and 5E , one of the auxiliary three-phase loads connected to output connections  340   a - c  of the PPSNDs  302   a - c  would be the standby MEC  48 . In such case, the tie  270  providing three-phase power from the standby MEC  48  would be a four wire conductor with a separate wire connected to each of the three PPSNDs  302   a - c  and the fourth wire as a neutral connected to a fourth PPSND  302 . Although  FIGS. 12A-12C  depict direct connections for a total of two different three-phase loads, many other three-phase loads may be serviced by a particular primary MEC  44  with additional connections. 
     One or more of the MECS  44 ,  46 ,  48  may also include an integrated truss system having a mounting structure of one or more data and/or power transfer layers separated by one or more insulation layers. The truss is configured to facilitate easy installation or replacement within an aircraft  10  and may be constructed of rigid or flexible materials such as sheet metal, thermoplastics, composites, or some other suitable material. In an aircraft, power or data could be transferred to various locations on the mounting structure of the truss system or to various locations in the aircraft. In some configurations, a via or a mechanism such as a truss interconnect can electrically connect one or more power or data lines in one layer to one or more power or data lines in one or more different layers of the integrated truss system, as described in U.S. patent application Ser. No. 13/930,024, entitled TRUSS INTERCONNECT, filed 28 Jun. 2013, which is incorporated herein by reference in its entirety. The interconnect can also be used to electrically interconnect a LRU mounted to the top surface layer of the integrated truss system and to send power into the truss or from the truss into the LRU. An LRU with the PPSNDs  302   a - c  has a conductive boss (projection) and as the interconnect passes through the LRU and into the truss system the interconnect expands into the boss as well as the transfer layers of the truss system to make electrical connections between the LRU and the truss system. 
     In some configurations, the integrated truss system may electrically connect both power and data systems. In further configurations, the truss interconnect can also provide a mechanical connection between one or more layers of the integrated truss system. In additional configurations, the truss interconnect may be configured for multiple insertions and extractions, allowing the reuse of the truss interconnect. 
       FIG. 13  illustrates an exploded, perspective view of a multi-layered integrated truss system  500  of a MEC  44 ,  46 ,  48 . The integrated truss system  500  may include insulation layers  502   a - 502   b  (hereinafter referred to collectively and/or generically as “insulation layers  502 ”) and transfer layers  504   a - 504   c  (hereinafter referred to collectively and/or generically as “transfer layers  504 ”). In some configurations, the insulation layers  502  and the transfer  504  layers are alternately arranged among each other such that the insulation layers at least partially electrically separate the transfer layers  504  from one another. In further configurations, the insulation layers  502  are configured to, at least partially, physically separate one or more of the transfer layers  504  from one or more other transfer layers  504 . Also, in some configurations one or more of the insulation layers may act as a smoke or water drip barrier between the passenger and cargo compartments. 
     Components of a MEC  44 ,  46 ,  48  may be detachably secured to the truss system  500 . A portion of the power busing network system  90  of  FIG. 3 , for example corresponding with the high power portion  120  of a primary MEC  44 , with the PPSNDs  302   a - c , is housed in an LRU  52  mounted to the top surface insulation layer  502   a  of the truss system  500 . Also inside the LRU  52  with the power busing network system  90  is a microprocessor that receives channel A and B data inputs from the CNI module  162  to control all the contactors  216 ,  218 ,  220 ,  222 ,  232 ,  250 ,  260  and  278 . 
     Three-phase primary power  506   a - d  (hereinafter may be referred to collectively and/or generically as “three-phase primary power  506 ”) is provided from one of the main generators  34 ,  36  to the PPSNDs  302   a - c  inside the power busing network system  90 . Phase A power  506   a , phase B power  506   b , or phase C power  506   c , or all three, may be routed from the output connections  390   a - c  through the insulation layers  502  to one or more transfer layers  504  of the truss system  500 . The neutral  506   d  of the three phase primary power  506  also may be routed through the insulation layers  502  to one or more transfer layers  504  of the truss system  500 . Communication data is sent from one MEC  44 ,  46 ,  48  to any other MEC  44 ,  46 ,  48  across two data channels  188 ,  190  (commonly referred to as channels A and B). As shown in  FIG. 13 , the mounting structure of the truss system  500  provides separate layers configured to provide separate communication channels to system components mounted to the truss system  500 . Both data channels  188 ,  190  may be routed through the insulation layers  502  to one or more transfer layers  504  of the truss system  500 . For example, the transfer layer  504   a  includes data transfer path  536  and transfer layer  504   b  includes data transfer path  538 . The data transfer paths  536 ,  538  may be separated from one another by one or more layers  502 ,  504  such as transfer layer  504   c . Data communications back and forth between the power busing network system  90  with PPSNDs  302  and the CNI module  162  are sent back and forth across the data channels  188 ,  190 . Data channel  188  passes through the transfer path  536  of transfer layer  504   a  and data channel  190  passes through the transfer path  538  of transfer layer  504   b.    
     In some configurations, the transfer layers  504  are configured to include one or more power or data transfer paths, or both. For example, the transfer layer  504   c  may include power transfer paths  512   a  and  512   b  which correspond with phase B power  506   b  and neutral  506   d  of the three phase primary power  506 . The power transfer path  512   a  receives phase B power, of 230 VAC for example, and transfers it to another LRU  52  mounted to the truss system  500  such as the CNI module  162  shown in  FIG. 13 . Transfer path  512   b  is the current return path across the neutral  506   d  from the CNI module  162  back to one of the PPSNDs  302 . 
     Each MEC  44 ,  46 ,  48  also includes at least one power distribution module  170  for distributing secondary power from the MECs  44 ,  46 ,  48 . Each distribution module  170  may be configured as one or more LRUs  52 . Each distribution module  170  preferably receives all three phases but distributes them to single phase leads in a balanced manner. As shown in  FIG. 13 , phase A  506   a  and phase B  506   b  are provided through two different transfer layers  504  of the truss system  500 , and then to distribution modes  170 . Each distribution module  170  then distributes single phase secondary power to the low power equipment loads  50  within the assigned zone of each particular MEC  44 ,  46 ,  48 . The equipment loads  50  associated with each MEC  44 ,  46 ,  48  are preferably distributed evenly across all three power phases. Preferably, each of the low power equipment loads  50  is connected to a distribution module  170  with a twisted electrical conductor pair. Although the present application depicts a particular number of connections in one or more of the Figures, any number of equipment loads  50  may be serviced by a MEC  44 ,  46 ,  48  subject to the amount of secondary power available. 
       FIG. 14  illustrates a primary MEC  44  with several layers of a truss system. The primary MEC  44  includes TRUs  134 ,  142 , the ATU  138 , the CNI module  162 , distribution modules  170  and a PPSND  302 . Primary MECs  44  include a PPSND  302  and MECs  46 ,  48  do not. A secondary MEC  46  could be depicted in a manner similar to the MEC  44  in  FIG. 10B  except without a PPSND  302 . The two TRUs  134 ,  142 , the ATU  138 , the CNI module  162 , distribution modules  170  and the PPSND  302  are electrically interconnected to the traces or metalized interconnects in the transfer layers  504  by inserting interconnector mechanisms  562 . The interconnection mechanisms  562  are inserted through each of the TRUs  134 ,  142 , the ATU  138 , the CNI module  162 , distribution modules  170  and the PPSND  302  and into vias  566  in each of the transfer layers  504 . 
     The truss system includes transfer layer  504   a  with trace  536  for channel A and transfer layer  504   b  with trace  538  for channel B. Each of the TRUs  134 ,  142 , the ATU  138 , distribution modules  170  and the PPSND  302  are connected to a dedicated channel A trace  536  and to a dedicated channel B trace  538 . However, the number of traces,  536 ,  538  on each transfer layer  504  depends on the protocol. In other embodiments, the TRUs  134 ,  142 , the ATU  138 , the distribution modules  170  and the PPSND  302  could all be connected to the same channel A trace  536  and to the same channel B trace  538 . 
     The truss system in  FIG. 14  also includes transfer layers  504   c ,  504   d ,  504   e  and  504   f . Transfer layer  504   c  includes traces  570  with three-phase primary power  506 , such as 230 VAC, for powering the truss system of the MEC  44 ,  46 ,  48  and the systems connected to it. A respective trace  570  corresponds with phase A power  506   a , phase B power  506   b , phase C power  506   c , and neutral  506   d . The two TRUs  134 ,  142 , the ATU  138 , the CNI module  162 , and the PPSND  302  are connected to the traces  570  of transfer layer  504   c  with interconnector mechanisms  562  through vias  566 . The three-phase primary power  506  is provided from the generators  34 ,  36  through the PPSND  302  to the transfer layer  504   c . The two TRUs  134 ,  142 , the ATU  138 , and the CNI module  162  are then powered by receiving the three-phase primary power  506  from the traces  570  of the transfer layer  504   c.    
     Secondary power is distributed from the TRUs  134 ,  142  and the ATU  138  to transfer layers  504   d ,  504   e ,  504   f . Transfer layers  504   d ,  504   e  are low voltage layers, such as 28 VDC, and each includes a positive trace  574 , a negative trace  576 , and a neutral trace  578 . One of these transfer layers  504 , such as transfer layer  504   e , may provide standby power from the fuel cell or RAT  128  via the second TRU  142 . 28 VDC power from the traces  574 ,  576 ,  578  of transfer layers  504   d ,  504   e  is distributed to the distribution module  170 . Transfer layer  504   f  is a low voltage three-phase layer, such as 115 VAC, that includes phase A power  580 , phase B power  582 , phase C power  584  and a neutral  586 . 115 VAC power from the traces of transfer layer  504   f  is also distributed to the distribution module  170 . 
     The distribution module  170  is connected to the traces of transfer layers  504   d ,  504   e ,  504   f  for secondary power and also to the traces  536 ,  538  for channels A and B  202 ,  204  in order to distribute the secondary power to equipment loads  50  with the twisted and shield electrical conductor pairs  314 . The distribution module  170  is not connected to transfer layer  504   b  with three-phase primary power  506  because primary power is not distributed from the distribution module  170 . Communication data from channels A and B  202 ,  204  of truss transfer layers  504   a ,  504   b  controls when the distribution module  170  turns on and off secondary power to the twisted and shield electrical conductor pairs  314  to service the equipment loads  50 . 
     As shown in  FIG. 14 , the CNI module  162  is connected to every trace in every layer  504  of the truss system of the MEC  44 ,  46 ,  48 . Because there are multiple voltage inputs to the CNI module  162 , power regulators perform conversions to the needed voltages. If any of the traces on one or more of the layers  504  become powered, the CNI module  162  wakes up. For example, if all the MECs  44 ,  46  lose primary power, power could be provided to the standby MEC  48  with a fuel cell or the RAT  128  thereby providing power to traces  574 ,  576 ,  578  of the standby layer  504   e . Power in the traces  574 ,  576 ,  578  of the transfer layer  504   e  would awaken the CNI module  162 . The CNI module  162  also receives communication data for use with the network switches  182 ,  184 ,  186  from both channels A and B  202 ,  204  from each of the traces  536 ,  538  of the transfer layers  504   a ,  504   b.    
       FIG. 14  also depicts a barrier  588  preferably positioned above the transfer layers  504  of the truss system of a MEC  44 ,  46 ,  48 . If the truss system were positioned within the floor structure as shown in  FIG. 17 , the barrier serves as a smoke barrier for obstructing smoke from the cargo compartment from entering the passenger compartment and/or as a water drip barrier for obstructing dripping water anywhere with the aircraft  10 . For example, the barrier  588  could obstruct water from dripping onto electrical components of a MEC  44 ,  46 ,  48 . Alternatively, or in addition to the barrier  588 , one or more of the insulation layers  502  could be the smoke and/or water drip barrier. For example, the uppermost insulation layer  502  of the truss system could be configured to act as a barrier to water and smoke. 
     In existing composite aircraft, the current return network provides a fault current return path, a personal safety protection path, and lightning protection path for the aircraft systems. However, as explained above, the current return network also provides a significant amount of wire weight to an aircraft which is undesirable. 
     The current return networks of these known aircraft are also susceptible to large voltage offsets. Both AC and DC voltages may be measured on the current return network. The return currents of all the equipment loads throughout the aircraft on the current return network are cumulative and therefore a voltage drop is created along the current return network as measured from the power source grounding point to the load grounding points. The voltage drop at different points along the current return network increases from the power source grounding points toward the back of the aircraft proportional to the impedance of the current return network and the current passing through it. 
       FIG. 15  generally illustrates three-phase (3Φ) primary power  506  is routed from one or more of the main generators  34 ,  36  to multiple isolated TRUs  134  and non-isolated ATUs  138 . The TRUs  134  and ATUs  138  are distributed throughout the aircraft  10  as part of the distributed architecture as shown in  FIG. 15 . At least one TRU  134  and at least one ATU  138  corresponds with one of the MECs  44 ,  46 ,  48 . Because the TRUs  134  are isolated they can be grounded wherever it is convenient. Also, because the TRUs  134  are distributed, the TRUs  134  can be grounded at different locations and, therefore, their DC return currents remain local to each respective MEC  44 ,  46 ,  48 . However, the return currents are no longer cumulative which results in a DC offset voltage of zero. 
       FIG. 16  also generally illustrates the distribution of either AC or DC power from an ATU  138  or a TRU  134 , respectively. However, more specifically as described above, the primary power  506  is first distributed to the power conversion equipment and then to the distribution modules  170  connected to each of the low power equipment loads  50  with multiple twisted and shielded electrical conductors where the conductors carry essentially equal but opposite currents. In application there may be small differences in current carried by the conductors. For example, twisted and shielded electrical conductor pair  314  includes an electrical power conductor  310  and a neutral or return conductor  312 . The neutral conductor may be routed with a three-phase power feeder. 
     After converting the primary power  506 , AC power is distributed from each ATU  138  to AC equipment loads  50   a  with an electrical power conductor  310  and current is returned from each AC equipment load  50   a  on a corresponding return conductor  312  of the twisted and shielded electrical conductor pair  314 . DC power is provided from each TRU  134  to the DC equipment loads  50   b  with electrical power conductor  310 . Current is returned from each DC equipment load  50   b  on the corresponding return conductor  312  of the twisted electrical conductor pair. 
     Phase A power  506   a , phase B power  506   b , and phase C power  506   c  are distributed from the generators  34 ,  36 . A fourth wire from the generators  34 ,  36  for the three-phase primary power  506  is also depicted that is the neutral conductor  506   d . Each of the AC equipment loads  50   a  includes a shield termination wire  590  depicted by a broken line connected to the neutral conductor  506   d  and each of the DC equipment loads  50   b  includes a shield termination wire  592  also depicted by a broken line connected to the neutral conductor  506   d . Although each of the equipment loads  50   a  and  50   b  are connected to the neutral conductor  506   d  with shield termination wires  590  and  592 , respectively, the load return currents are no longer cumulative. In  FIG. 16 , part of the neutral conductor  506   d  is configured to appear as a current return network (CRN) merely to illustrate that the voltage differential is zero as a result of using small loops of twisted wire conductor pairs for localized secondary power distribution. The neutral conductor  506   d  of the distributed three-phase primary power  506  between MECs  44 ,  46 ,  48  of the aircraft  10 , which is much smaller than conductors that would typically be utilized as part of a CRN, may simply be referred to as a safety ground bus (SGB). Therefore, a CRN is no longer needed in the composite aircraft  10  with localized secondary power distribution provided by twisted wire conductor pairs. The twisted wire conductor pair now provides current return. Also, the cross-sectional area of the loops created by the twisted conductor pair is much smaller than the cross-sectional area created by the larger wire loop of the CRN which reduces the lighting threat to the composite aircraft  10 . For comparison, the conductors of the twisted pair may be about 16 to about 20 American wire gauge (AWG) whereas the conductors of the CRN are about 2 AWG or larger diameter. 
       FIG. 16  also illustrates the distribution of primary power from generators  34 ,  36  among primary MECs  44  distributed within the forward, mid and aft sections of the aircraft  10 . Each primary MEC  44  includes a TRU  134  and an ATU  138  for servicing equipment loads  50   b  and equipment loads  50   a , respectively, as described above. Power is distributed from each MEC  44  to each equipment load  50  with a twisted and shielded electrical conductor pair  314 .  FIG. 16  also depicts a pair of MECs  44  providing 230 VAC for auxiliary loads  520 . As referenced in  FIGS. 12A-12C  and the accompanying text, 230 VAC power to and from the auxiliary loads is controlled by the contactors  232 ,  278  of the PPSNDs  302  of the primary MEC  44 . 
       FIG. 16  also illustrates a plurality of LRUs  52 , such as avionics, serviced by the forward most primary MEC  44 .  FIG. 16  also illustrates a battery  598  for providing standby power. Although  FIG. 16  depicts the battery  598  providing standby power to only the forward most primary MEC  44 , battery standby power is preferably provided to all primary MECs  44 . 
       FIG. 17  illustrates an integrated truss system  600  which may be used in aircraft manufacturing for providing one or more power and data transfer paths as explained above. One or more MECs  44 ,  46 ,  48  may include the truss system  600  as a support or mounting structure for attaching all or part of vehicle systems, components of a MEC  44 ,  46 ,  48 , equipment loads  50 , LRUs  52 , or other equipment. 
     The mounting structure of the truss system  600  may be a multi-part or modular assembly of separate structural elements that stack, detachably connect or lock together to create an integrated mounting structure that may be installed in an aircraft  10  as a single unitary piece. Each structural element may have one or more transfer layers and one or more insulation layers as described above. Each structural element of the multi-part truss system  600  may be detachable from one another to allow repair or replacement of damaged structural elements without removing undamaged structural elements from the aircraft  10 . One or more layers of each structural element may also be replaced. One element of the truss system  600  could be swapped out without having to remove the entire truss system  600 . Also, all or at least a portion of the truss system  600  may also be detachable from the support structure of the aircraft  10  such as the floor beams or fuselage frame members. Alternatively, the truss system  600  may be manufactured as a single monolithic structure which may be installed or replaced in it&#39;s entirely. 
     The truss system  600  is configured to extend within a thin structural volume defined in the sidewall of the fuselage between frame members, and by the depth of the frame members, or in the space in the floor between the passenger and cargo compartments of the aircraft  10 , and by the depth of the floor beams. Alternately, the truss system  600  could have a physical form configured to be implemented within a traditional equipment bay. The truss system  600  mounted in the sidewall of the aircraft  10  preferably corresponds with the curvature of the fuselage of the aircraft  10 .  FIG. 17  is a bottom view looking upward toward the truss system  600  configured to extend from sidewall to sidewall of the aircraft  10 , under seat rails  610 , and between transverse floor beams  608 . A MEC  44 ,  46 ,  48  positioned in the floor or in the sidewall of the aircraft  10  with a truss system such as truss system  600  can service the equipment loads  50  within the passenger compartment and in the cargo compartment of the aircraft  10  that are in proximity of the MEC  44 ,  46 ,  48 . 
     The truss system  600  is configured to have a narrow middle portion that extends over the top of two inner adjacent floor beams  608  and opposite end portions that extend further outward from both sides of the two inner adjacent floor beam  608  to the next floor beams  608  to provide a wide surface for mounting components such as the power distribution modules  170 . In one or more embodiments, the truss system is configured to have a width and length between adjacent floor beams  608 , or between floor beams  608  that are displaced from one another, that is suitable for serving as a smoke barrier for obstructing smoke from the cargo compartment from entering the passenger compartment and/or as a water drip barrier for obstructing water dripping onto electrical components within the MEC  44 ,  46 ,  48 . 
       FIG. 17  also shows the CNI module  162 , power distribution modules  170 , TRUs  134 ,  142 , the ATU  138 , and the PPSNDs  302  mounted to the truss system  600  of a primary MEC  44 . The TRU  134  receives 230 VAC from the output connections  390  of the PPSNDs  302 . The TRUs  134  connect to a power bus with 28 VDC to power the distribution modules  170 . Each power distribution module  170  has connections  596  for interfacing with the equipment loads  50  associated with the primary MEC  44 . 
     Each structural element of the truss system  600  has one or more transfer and insulation layers as explained above. One of the transfer layers may be configured to transfer high voltage power from one portion of a MEC  44 ,  46 ,  48  to another portion of that same MEC  44 ,  46 ,  48 . For example, high voltage power may be provided inside the truss system  600  across a transfer layer to the PPSNDs  302 , configured as an LRU  52 , mounted to the surface of truss system  600 . Low voltage secondary power may also be provided through another transfer layer of the truss system  600  to low power equipment loads  50  mounted to the surface of the truss system  600 . Also, communication data can be provided across a transfer layer of the truss  600  to an aircraft system component mounted to the surface of the truss system  600 . One transfer layer of the truss system  600  could provide channel A to a system component mounted to the surface of the truss system  600  and another transfer layer could provide channel B to that same system. 
     Often electronic protective functions are implemented only after the design of an aircraft is complete which lends itself to longer response times because of unintended latencies. One or more configurations of the aircraft  10  may have distributed electronic protections and a control architecture that enables multiple fault clearing actions to occur simultaneously without coordination issues due to conflicting controllers involved in fault isolation logic. A universal or common power controller such as an integrated protection chipset (IPC) scheme within the aircraft  10  can be easily scaled up or down based on the desired configurations. Also, a universal IPC allows for scaling of performance and the extension of deterministic protection behavior in a localized way. Parallel processing by the IPCs within the distributed architecture of the aircraft  10 , where each IPC performs independent protective management, allows faster processing of fault clearances and corrective actions of electrical faults internal to one or more MECs  44 ,  46 ,  48  or on wiring downstream of the MECs  44 ,  46 ,  48 . A cycle of commands for control, protection and safeguarding electrical systems, generators, and feeders occurs below a voltage cycle of about 1.25 milliseconds for a 800 Hz frequency. 
     Protective functions such as, but not limited to, overcurrent protection such as I 2 t (ampere-squared-seconds), differential protection, ground fault protection, corona fault detection, arc fault detection, overvoltage protection, undercurrent protection, voltage sensing, current sensing, current transfer, over temperature protection, open phase protection, unbalanced current protection, voltage RMS/phase functionality and impedance assessment using reflectometry methods are blended together and integrated onto a single IPC to reduce costs and complexity and to minimize supplier implementation differences that erode protection coordination performance. Trip values and specific protection functionality are customized at each location through data load, pin program or bar code reader in conjunction with lookup tables. Also, determining distance to a fault location via reflectometry impedance assessments may be integrated into the IPCs without association with electrical protective functions. 
     Traditionally, a controller sensing what is going on with one or more contactors sees heavy utilization sampling and broadcasting data, only to wait for an outside entity to make a determination on its state. This introduces latency into the communication network. Each IPC can make its own independent fault assessments and perform protective functions locally without dependence on internal data buses or isolation logic performed by multiple other power controllers distributed throughout the aircraft communicating with one another to perform fault isolation logic. For example, IPCs within a MEC  44 ,  46 ,  48  have knowledge of the secondary power distributed downstream to the associated equipment loads  50 . These IPCs within a MEC  44 ,  46 ,  48  can locally sense any faults, perform local protective functions, and interrupt accordingly without also requiring knowledge of and independent from what is occurring upstream in the aircraft  10  such as on the primary power switching network or with the controllers of the power sources that provide power to the MECs  44 ,  46 ,  48 . 
     IPCs may be utilized anywhere power protections are needed. For example, IPCs may be used for interrupting fault currents anywhere there are contactors. Each contactor can act autonomously but the status of its own protective functions is communicated to each of the other IPCs in order to configure themselves based on data originating somewhere adjacent to them and received across the communication network. For example, a status of protective functions implemented by one IPC of one MEC is communicated to one or more IPCs of one or more other MECs. Because each contactor can independently determine and initiate its own protection, the faults occurring in local areas are cleared faster and there is less communication network overhead. Thus, the processing of fault logic is pushed outward and performed in parallel at the perimeter of the power distribution system as opposed to at the center. Otherwise, centralized processing takes up a lot of processor resources. 
       FIG. 18  illustrates IPCs  824  where power feeders  40 ,  42 ,  130  interconnect with the main power generators  34 ,  36 , the auxiliary power unit generator  54 , and the MECs  44 ,  46 ,  48 . A generator controller such as generator control unit (GCU)  820  for the generator  34 ,  36  and for the auxiliary power unit generator  54  each includes an IPC  824 . Thus, the GCU  820  with IPC  824  is delegated to contain high power level protections that independently prevent damage to the generators  34 ,  36  and power feeders  40 ,  42  along with power quality protections and to contain protections that independently prevent damage to the auxiliary power unit generator  54  and power feeder  130  along with power quality protections. 
     The IPCs  824  on-board a MEC  44 ,  46 ,  48  contain high power level protections that can perform coordinated assessments with the IPCs  824  of other MEC  44 ,  46 ,  48  that prevent damage to the associated MEC  44 ,  46 ,  48 , the corresponding power feeders  40 ,  42 ,  130 , as well as the corresponding forward, aft, and mid tie buses  76 ,  78 ,  80 . The IPCs  824  residing where primary power is provided to the MECs  44 ,  46 ,  48  perform high power level protection actions in conjunction with IPCs  824  of the controller units associated with the primary power sources such as the GCU  820  or across the forward, aft or mid tie buses  76 ,  78 ,  80  providing primary power from adjacent MECs  44 ,  46 ,  48 . For example, where there is a flow of current between MECs  44 ,  46 ,  48 , an IPC  824  requires knowledge of what is going on with the other MECs  44 ,  46 ,  48  as well as with its own MEC to coordinate the high power level protective functions. Also, IPCs  824  of the MECs  44 ,  46 ,  48  need to know what is going on with the IPCs  824  of the generators  34 ,  36  and vice versa. 
     Other IPCs  824  on-board a MEC  44 ,  46 ,  48  also provide zonal independent low power level protections where each IPC  824  performs assessments independent of other power controllers, such as independent of power controllers of the power source, for the zonal distribution of secondary power across twisted and shielded electrical conductor pairs  314  to the equipment loads  50  within the zone associated with each MEC  44 ,  46 ,  48 . By implementing an IPC  824  with each contactor, protections may be implemented for controlling a specific equipment load  50 . Because power and quality protections can be performed locally by each one of the on-board IPCs  824  and independent from centralized power controllers such as the GCUs  820  or a bus power control unit, the control logic conflicts and the high volume of communications that result in delays in clearing faults associated with the secondary power distribution are eliminated. Thus, the IPCs  824  allow the universal coordination of inter-MEC type high power level protections between the generators  34 ,  36 , the auxiliary power unit  56  and the MECs  44 ,  46 ,  48  while at the same time also having lower power level protections for the equipment loads  50  distributed throughout the aircraft  10 . 
     Protection schemes of one or more IPCs  824  may be defined in terms of a zone. For example, two or more IPCs  824  may define a zone of protection where each one of the IPCs  824  is response for a portion of the zone and their assessments are coordinated to locate and locally isolate faults within that zone. Also, IPCs  824  along a perimeter of a zone act collectively to provide fault assessments for their zone of protective functions. Specific protection zones could be defined based on a specific architecture such as, but not limited to, a tie bus zone, a primary power switching bus zone, a zone of a single MEC, a zone of multiple MECs, a primary MEC  44  to primary MEC  44  zone, a generator to one or more MECs zone, or a zone of equipment loads associated with a MEC  44 ,  46 ,  48 . IPCs  824  such as those performing high power level protection where primary power is provided from the generators  34 ,  36  to the MECs  44 ,  46 ,  48  may be defined as a zone of high power level protection. Also, a standalone IPC  824  that provides independent fault assessments could be its own zone. For example, between an input to an output within a truss of a MEC  44 ,  46 ,  48  may be a zone for a standalone IPC  824 . Also, IPCs  824  performing low power level protections for downstream secondary power distribution from the MECs  44 ,  46 ,  48  across twisted and shielded electric conductor pairs  314  may be defined as a zone of low level protection. 
     For differential current protection, currents are monitored at multiple locations in a zone. If a current mismatch exists in the zone, specific contactors would be opened to isolate the fault. If there are failures in the power distribution system, such as a PPSND  302  loses communications or is otherwise inoperable requiring power routing reconfiguration, the zone that is being protected could be adapted or expanded to include additional current sensing in other locations such as in other MECs and different contactors could open to isolate a fault. Also, if an IPC  824  fails within a zone, that zone may be adapted or expanded to include one or more additional IPCs  824  to maintain a perimeter that is protectively functional. 
       FIG. 19  illustrates one configuration of one of the MECs  44 ,  46 ,  48  having a distribution module  170  and a PPSND  302 . However, the MEC  44 ,  46 ,  48  of  FIG. 19  includes other components as described in more detail herein. The distribution module  170  and the PPSND  302  each include multiple contactors such as the contactors  842  shown in  FIG. 19 . The contactors  842  correspond to any one of the contactors depicted or described herein in association with the distribution modules  170 , PPSNDs  302 , any other device of the MECs  44 ,  46 ,  48 , or elsewhere on the aircraft  10 , where power protection is needed. For example, at least one configuration of each PPSND  302  includes eight contactors  210 ,  218 ,  220 ,  222 ,  232 ,  250 ,  260  and  278 , as described above, where each functionally corresponds with a contactor  842  having its own protections through a corresponding IPC  824 . 
     Multiple switching elements within each of the contactors  842  are each commanded to open and close by gate drive electronics for controlling the contactors  842 . The IPCs  824  are associated with the gates drive electronics as wells as current sensing electronics for monitoring the current through each of the switching elements. If the electronics detect some type of fault such as a ground or arc fault, overvoltage, overcurrent, undercurrent, etc., the gate of one or more switching elements is commanded to open to interrupt the current through the contactor  842 . Each IPC  824  includes a DSPic or microprocessor  826  and an application-specific integrated circuit (ASIC)  828  customized for locating and identifying faults. All the analog and digital sensed data is pre-processed and digitized via the ASIC  828  allowing the microprocessor  826  to be programmed with universal, common protection software with the desired level of design assurances. This digitized data is available to the CNI module  162  of the MEC  44 ,  46 ,  48  to communicate with other CNI modules  162  and IPCs  824  within other MECs  44 ,  46 ,  48  within adjacent zones. For example, the IPC  824  could include spread-spectrum time-domain reflectometry (SSTDR) for locating and identifying faults on powered circuits. SSTDR looks across a broader frequency range on a powered circuit that enables reconstructions of the reflections back to obtain information about any circuit faults on the powered circuit. 
     Each distribution module  170  could have a dedicated IPC  824  installed in each of its contactors or a single IPC  824  could be multiplexed across multiple contactors depending on the speed of the IPC  824 &#39;s microprocessor  826 . Also, each PPSND  302  could have a dedicated IPC  824  for each of its contactors or a single IPC  824  could be multiplexed across multiple contactors depending on the speed of the IPC  824 &#39;s microprocessor  826 . 
     Digital differential protection integrated into the IPCs  824  monitors current on both ends of each wire of the primary power distribution network with current sensing integrated into each IPC  824  and therefore located at the generators  34 ,  36  and at the MECs  44 ,  46 ,  48  for transferring power from one MEC  44 ,  46 ,  48  to another. The current sensing look at the current between the two different locations of the two wire ends that are spaced some distance apart and then share the data digitally with other IPCs  824 . Separate external differential protection circuitry is not needed because the differential protection is integrated into the solid state contactors and associated controls which reduce the overall wiring complexity and weight of the aircraft  10 . 
     Data is digitized at each current sensor and then sent over the communications network to do the current comparisons and ascertain whether there is a fault somewhere in the power distribution system. If the current at two locations is different or the sum of currents downstream of the source are different than the source that exceeds a set threshold then there is a fault. For example, on a PPSND  302  the contactor  220  at the output connection  214  for a secondary MEC  46  has an associated IPC  824 . The current (magnitude and relative phase angle) is measured at the contactor  220  and at the IPC  824  at the power input to the secondary MEC  46 . One or more communication channels exist between a primary MEC  44  and a secondary MEC  46 . The IPCs  824  distributed amongst the MECs  44 ,  46 ,  48  and the GCUs  820  compare the contactor currents to ascertain whether faults exits and where. Also, the use of the IPCs  824  with the contactors enables the classification of the fault type to enable adaptive protection action or system isolation, distinguishing between fault currents and magnetization currents to avoid unwanted protective actions, and provide built-in test equipment data for localizing the fault for repair or replacement thereby reducing repair time. 
     On the primary MEC  44  the current flow out of the contactor  220  to the secondary MEC  46  is compared to the incoming current flow to the secondary MEC  46  and the digital differential protection is able to communicate fast enough across the existing communication network between primary and secondary MECs  44 ,  46  to time the two current readings to make reliable fault current determinations. Thus, each IPC  824  makes its own non-centralized control determination for different fault assessments based upon the shared data provided by network switches  182 ,  184 ,  186  of other MECs  44 ,  46 ,  48 . The distributed MEC architecture with IPCs  824  of the aircraft  10  allows for localization of protective actions that would otherwise have been performed or coordinated by other elements that introduce latencies to response times. 
     Turning now to  FIG. 20 , an illustrative routine  900  for providing distributed protection and control architecture for electrical systems of a vehicle is provided herein. Unless otherwise indicated, more or fewer operations may be performed than shown in the figures and described herein. Additionally, unless otherwise indicated, these operations may also be performed in a different order than those described herein. 
     The routine  900  starts at operation  902 , where MECs  44 ,  46 ,  48  are spatially distributed throughout a vehicle. In operation  904  primary power is distributed to the MECs  44 ,  46 ,  48 . Operation  906  includes distributing secondary power from each MEC  44 ,  46 ,  48  to equipment loads  50  nearest each MEC  44 ,  46 ,  48 . Operation  908  includes providing IPCs  824  having electronic protective functions. In operation  910  the IPCs  824  perform high power level protections by performing coordinated fault assessments where primary power is received at each MEC  44 ,  46 ,  48 . In operation  912  the IPCs  824  independently perform low power level protections where secondary power is distributed from each MEC  44 ,  46 ,  48 . 
     The subject matter described above is provided by way of illustration only and should not be construed as limiting. Various modifications and changes may be made to the subject matter described herein without following the example embodiments and applications illustrated and described, and without departing from the true spirit and scope of the present disclosure, which is set forth in the following claim.