Patent ID: 12246607

DETAILED DESCRIPTION

System Overview

FIG.1Aillustrates an aircraft100, such as an electric or hybrid aircraft, andFIG.1Billustrates a simplified block diagram of the aircraft100. The aircraft100includes a motor110, a management system120, and a power source130. The motor110can be used to propel the aircraft100and cause the aircraft100to fly and navigate. The management system120can control and monitor the components (for example, equipment) of the aircraft100, such as the motor110and the power source130. The power source130can power the motor110to drive the aircraft100and power the management system120to enable operations of the management system120. The management system120can include one or more motor controllers as well as other electronic circuitry for controlling and monitoring various components of the aircraft100.

FIG.2illustrates components200of an aircraft, such as the aircraft100ofFIGS.1A and1B. The components200can include a power management system210, a motor management system220, and a recorder230, as well as a first battery pack212A, a second battery pack212B, a warning panel214, a fuse and relay216, a converter217, a cockpit battery pack218, a motor controller222, one or more motors224, and a throttle226.

The power management system210, the motor management system220, and the recorder230can monitor communications on a communication bus, such as a controller area network (CAN) bus, and communicate via the communication bus. The first battery pack212A and the second battery pack212B can, for instance, communicate on the communication bus enabling the power management system210to monitor and control the first battery pack212A and the second battery pack212B. As another example, the motor controller222can communicate on the communication bus enabling the motor management system220to monitor and control the motor controller222.

The recorder230can store some or all data communicated (such as component status, temperature, or over/undervoltage information from the components or other sensors) on the communication bus to a memory device for later reference, such as for reference by the power management system210or the motor management system220or for use in troubleshooting or debugging by a maintenance worker. The power management system210and the motor management system220can each output or include a user interface that presents status information and permits system configurations. The power management system210can control a charging process (for instance, a charge timing, current level, or voltage level) for the aircraft when the aircraft is coupled to an external power source to charge a power source of the aircraft, such as the first battery pack212A or the second battery pack212B.

The warning panel214can be a panel that alerts a pilot or another individual or computer to an issue, such as a problem associated with a power source like the first battery pack212A. The fuse and relay216can be associated with the first battery pack212A and the second battery pack212B and usable to transfer power through a converter217(for example, a DC-DC converter) to a cockpit battery pack218. The fuse and relay216can protect one or more battery poles of the first battery pack212A and the second battery pack212B from a short or overcurrent. The cockpit battery pack218may supply power for the communication bus.

The motor management system220can provide control commands to the motor controller222, which can in turn be used to operate the one or more motors224. The motor controller can include an inverter for generating AC currents that are needed for operating the one or more motors. The motor controller222may further operate according to instructions from the throttle226that may be controlled by a pilot of the aircraft. The one or more motors can include an electric brushless motor.

The power management system210and the motor management system220can execute the same or similar software instructions and may perform the same or similar functions as one another. The power management system210, however, may be primarily responsible for power management functions while the motor management system220may be secondarily responsible for the power management functions. Similarly, the motor management system220may be primarily responsible for motor management functions while the power management system210may be secondarily responsible for the motor management functions. The power management system210and the motor management system220can be assigned respective functions, for example, according to system configurations, such as one or more memory flags in memory that indicate a desired functionality. The power management system210and the motor management system220may include the same or similar computer hardware.

The power management system210can automatically perform the motor management functions when the motor management system220is not operational (such as in the event of a rebooting or failure of the motor management system220), and the motor management system220can automatically perform the power management functions when the power management system210is not operational (such as in the event of rebooting or failure of the power management system210). Moreover, the power management system210and the motor management system220can take over the functions from one another without communicating operation data, such as data about one or more of the components being controlled or monitored by the power management system210and the motor management system220. This can be because both the power management system210and the motor management system220may be consistently monitoring communications on the communication bus to generate control information, but the control information may be used if the power management system210and the motor management system220has primary responsibility but not if the power management system210and the motor management system220does not have primary responsibility. Additionally or alternatively, the power management system210and the motor management system220may access data stored by the recorder230to obtain information usable to take over primary responsibility.

System Architecture

Electric and hybrid aircraft (rather than aircraft powered during operation by combustion) have been designed and manufactured for decades. However, electric and hybrid aircraft have still not yet become widely used for most transport applications like carrying passengers or goods.

This failure to adopt may be in large part because designing an aircraft that is sufficiently safe to be certified by certification authorities may be very difficult. The certification of prototypes may moreover not be sufficient to certify for commercial applications. Instead, a certification of each individual aircraft and its components may be required.

This disclosure provides at least some approaches for constructing electric powered aircraft from components and systems that have been designed to pass certification requirements so that the aircraft itself may pass certification requirements and proceed to active commercial use.

Certification requirements can be related to a safety risk analysis. A condition that may occur with an aircraft or its components can be assigned to one of multiple safety risk assessments, which may in turn be associated with a particular certification standard. The condition can, for example, be catastrophic, hazardous, major, minor, or no safety effect. A catastrophic condition may be one that likely results in multiple fatalities or loss of the aircraft. A hazardous condition may reduce the capability of the aircraft or the operator ability to cope with adverse conditions to the extent that there would be a large reduction in safety margin or functional capability crew physical distress/excessive workload such that operators cannot be relied upon to perform required tasks accurately or completely or serious or fatal injury to small number of occupants of aircraft (except operators) or fatal injury to ground personnel or general public. A major condition can reduce the capability of the aircraft or the operators to cope with adverse operating condition to the extent that there would be a significant reduction in safety margin or functional capability, significant increase in operator workload, conditions impairing operator efficiency or creating significant discomfort physical distress to occupants of aircraft (except operator), which can include injuries, major occupational illness, major environmental damage, or major property damage. A minor condition may not significantly reduce system safety such that actions required by operators are well within their capabilities and may include a slight reduction in safety margin or functional capabilities, slight increase in workload such as routine flight plan changes, some physical discomfort to occupants or aircraft (except operators), minor occupational illness, minor environmental damage, or minor property damage. A no safety effect condition may be one that has not effect on safety.

An aircraft can be designed so that different monitoring and warning subsystems, such as battery monitoring circuits, of the aircraft are constructed to have a robustness corresponding to their responsibilities and any related certification standards, as well as potentially any subsystem redundancies.

Where a potential failure of the responsibilities of a monitoring and warning subsystem would likely be catastrophic, the subsystem can be designed to be simple and robust and thus may be able to satisfy difficult certification standards. The subsystem, for instance a battery, motor or motor controller monitoring circuit, can be composed of non-programmable, non-stateful components (for example, analog or non-programmable combinational logic electronic components) rather than programmable components (for example, a processor, a field programmable gate array (FPGA), or a complex programmable logic device (CPLD)) or stateful components (for example, sequential logic electronic components) and activate indicators such as lights rather than more sophisticated displays.

On the other hand, where either (i) a monitoring and warning subsystem (such as a battery monitoring circuit, a motor monitoring circuit or a motor controller monitoring circuit) of an aircraft monitors a parameter redundantly with another subsystem of the aircraft that is composed of non-programmable, non-stateful components or (ii) a potential failure of the responsibilities of such a monitoring and warning subsystem would likely be less than catastrophic, or less than hazardous, the subsystem can be at least partly digital and designed to be complicated, feature-rich, and easier to update and yet able to satisfy associated certification standards. Such a subsystem can, for instance, include a processor or other programmable components that outputs information to a sophisticated display for presentation.

In some implementations, some or all catastrophic conditions monitored for by an aircraft can be monitored for with at least one monitoring and warning subsystem that does not include a programmable component or a stateful component because certifications for programmable components or stateful components may demand statistical analysis of the responsible subsystems, which can be very expensive and complicated to certify. Such implementations can moreover be counterintuitive at least because an electric or hybrid aircraft may include one or more relatively advanced programmable or stateful components to enable operation of the electric or hybrid aircraft, so the inclusion of one or more subsystems in the aircraft that does not include any programmable components or any stateful components may be unexpected because the one or more relatively advanced programmable or stateful components may be readily and easily able to implement the functionality of the one or more subsystems that does not include any programmable components or any stateful components.

An aircraft monitoring system can include a first monitoring and warning subsystem and a second monitoring and warning subsystem. The second subsystem, such as a second battery monitoring circuit, can be supported by an aircraft housing and include non-programmable, non-stateful components, such as analog or non-programmable combinational logic electronic components. The non-programmable, non-stateful components can monitor a component (such as battery cells in a battery pack) supported by the aircraft housing and output a second alert to notify of a catastrophic condition associated with the component. The non-programmable, non-stateful components can, for instance, activate an indicator or an audible alarm for a passenger aboard the housing to output the first alert. The indicator or audible alarm may remain inactive unless the indicator is outputting the first alert. Additionally or alternatively, the non-programmable, non-stateful components can output the second alert to a computer aboard or remote from the aircraft (for example, to automatically trigger actions to attempt to respond to or address the catastrophic condition, such as to stop charging or activate a fire extinguisher, a parachute, or an emergency landing procedure or other emergency response feature) or an operator of the aircraft via a telemetry system. The non-programmable, non-stateful components may, moreover, not be able to control the component or at least control certain functionality of the component, such as to control a mode or trigger an operation of the component.

The first subsystem, such as a first battery monitoring circuit, can be supported by the aircraft housing and include a processor (or another programmable or stateful component), as well as a communication bus. The processor can monitor the component from communications on the communication bus and output a first alert to notify of a catastrophic condition or a less than catastrophic condition associated with the component. The processor can, for instance, activate an indicator or audible alarm for a passenger aboard the housing to output the first alert. Additionally or alternatively, the processor can output the first alert to a computer aboard or remote from the aircraft (for example, to automatically trigger actions to attempt to address the catastrophic condition, such as to activate a fire extinguisher, a parachute, or an emergency landing procedure) or an operator of the aircraft via a telemetry system. The processor may control the component.

The non-programmable, non-stateful components of the second subsystem additionally may not be able to communicate via the communication bus. It may not include any programmable communication circuit for allowing communication via such a bus.

An example of such a design and its benefits are next described in the context of battery management systems. Notably, the design can be additionally or alternatively applied to other systems of a vehicle that perform functions other than battery management, such as motor and motor control.

FIG.3illustrates a battery monitoring system. This system can be used in an electric vehicle, such as an electric aircraft, a large size drone or unmanned aerial vehicle, an electric car, or the like, to monitor the state of battery cells71in one of multiple battery packs and report this state or generate warning signals in case of dysfunctions.

The battery cells71can be connected in series or in parallel to deliver a desired voltage and current.FIG.3shows serially connected battery cells. The total number of battery cells71may exceed 100 cells in an electric aircraft. Each of the battery cells71can be made up of multiple elementary battery cells in parallel.

A first battery monitoring circuit can control and monitor the state of each battery cell71. The first battery management circuit can include multiple BMSs72, each of the BMSs72managing and controlling one of the battery cells71. The BMSs72can each be made up of an integrated circuit (for instance, a dedicated integrated circuit) mounted on one printed circuit board (PCB) of the PCBs90. One of the PCBs90can be used for each of the battery cells71or for a group of battery cells.FIG.4illustrates example components of one of the BMSs72.

The control of a battery cell can include control of its charging and discharge cycles, preventing a battery cell from operating outside its safe operating area, or balancing the charge between different cells.

The monitoring of one of the battery cells71by one of the BMSs72can include measuring parameters of the one of the battery cells71, to detect and report its condition and possible dysfunctions. The measurement of the parameters can be performed with battery cell parameter sensors, which can be integrated in the one of the BMSs72or connected to the one of the BMSs72. Examples of such parameter sensors can include a temperature sensor91, a voltage sensor92, or a current sensor. An analog-to-digital converter93can convert the analog values measured by one or more of the parameter sensors into multivalued digital values, for example, 8 or 16 bits digital parameter values. A microcontroller94, which can be part of each of the BMSs72, can compare the values with thresholds to detect when a battery cell temperature, battery cell voltage, or battery cell current is outside a range.

The BMSs72as slaves can be controlled by one of multiple first master circuits75. In the example ofFIG.3, each of the first master circuits75can control four of the BMSs72. Each of the first master circuits75can control eight of the BMSs72, or more than eight of the BMSs72. The first master circuits75can control more BMS and more battery cells in yet other implementations. The first master circuits75can be connected and communicate over a digital communication bus85.

The first master circuits75can also be connected to a computer79that collects the various digital signals and data sent by the first master circuits75, and may display information related to the battery state and warning signals on a display83, such as a matrix display. The display83may be mounted in the vehicle's cockpit to be visible by the vehicle's driver or pilot. Additionally or alternatively, the computer79can output the information to a computer remote from the aircraft or to control operations of one or more components of the aircraft as described herein.

The BMSs72can be connected to the first master circuits75over a digital communication bus, such as a CAN bus. A bus driver95can interface the microcontroller94with the digital communication bus and provide a first galvanic isolation89between the PCBs90and the first master circuits75. In one example, the bus drivers of adjacent BMSs72can be daisy chained. For example, as shown in FIG.4, the bus driver95is connected to the bus driver97of the previous BMS and to the bus driver98of the next BMS.

Each of the BMSs72and their associated microcontrollers can be rebooted by switching its power voltage Vcc. The interruption of Vcc can be controlled by the first master circuits75over the digital communication bus and a power source96.

FIG.3further illustrates a second battery monitoring circuit, which can be redundant of the first battery monitoring circuit. This second battery monitoring circuit may not manage the battery cells71; for example, the second battery monitoring circuit may not control charge or discharge cycles of the battery cells71. The function of the second battery monitoring circuit can instead be to provide a separate, redundant monitoring of each of the battery cells71in the battery packs, and to transmit those parameters or warning signals related to those parameters, such as to the pilot or driver or a computer aboard or remote from the aircraft as described herein. The second battery monitoring circuit can monitor the state of each of the battery cells71independently from the first battery monitoring circuit. The second battery monitoring circuit can include one of multiple cell monitoring circuits73for each of the battery cells. The parameters or warning signals may moreover, for example, be used by the second battery monitoring circuit to stop charging (for instance, by opening a relay to disconnect supply of power) of one or more battery cells when the one or more battery cells may be full of energy and a computer of the aircraft continues to charge the one or more battery cells.

Motor and Battery System

Battery packs including multiple battery cells, such as lithium-ion cells, can be used in electric cars, electric aircraft, and other electric self-powered vehicles. The battery cells can be connected in series or in parallel to deliver an appropriate voltage and current.

In electrically driven aircraft, the battery packs can be chosen to fulfil the electrical requirements for various flight modes. During short time periods like take off, the electrical motor can utilize a relatively high power. During most of the time, such as in the standard flight mode, the electrical motor can utilize a relatively lower power, but may consume a high energy for achieving long distances of travel. It can be difficult for a single battery to achieve these two power utilizations.

The use of two battery packs with different power or energy characteristics can optimize the use of the stored energy for different flight conditions. For example, a first battery pack can be used for standard flight situations, where high power output may not be demanded, but a high energy output may be demanded. A second battery pack can be used, alone or in addition to the first battery pack, for flight situations with high power output demands, such as take-off manoeuvring.

An electrical powering system can charge the second battery pack from the first battery pack. This can allow recharging of the second battery pack during the flight, subsequent to the second battery pack being used in a high power output demanding flight situation. Therefore, the second battery pack can be small, which can save space and weight. In addition, this can allow different battery packs for different flight situations that optimize the use of the battery packs.

The electrical powering system can also charge the second battery pack by at least one motor which works as generator (the motor may also accordingly be referred to as a transducer). This can allow recharging of the second battery pack during the flight or after the second battery pack has been used in a high power output demanding flight situation. Therefore, the second battery pack can be small, which can save space and weight. In addition, the different battery packs can allow the recovery of braking energy. Braking energy during landing or sinking recovered by a generator motor can create high currents which may not be recovered by battery packs used for traveling long distances. By using a second battery pack suitable for receiving high power output in a short time, more braking energy can be recovered via the second battery pack than the first battery pack, for example.

The electrical powering system can also include a third battery pack, which includes a supercapacitor. Because supercapacitors can receive and output large instantaneous power or high energy in a short duration of time, the third battery pack can further improve the electrical powering system in some instances. A supercapacitor may, for example, have a capacitance of 0.1 F, 0.5 F, 1 F, 5 F, 10 F, 50 F, 100 F, or greater or within a range defined by one of the preceding capacitance values.

Modular Battery System

The power sources in an electric or hybrid aircraft can be modular and distributed to optimize a weight distribution or select a center of gravity for the electric or hybrid aircraft, as well as maximize a use of space in the aircraft. Moreover, the batteries in an electric or hybrid aircraft can desirably be designed to be positioned in place of a combustion engine so that the aircraft can retain a similar shape or structure to a traditional combustion powered aircraft and yet may be powered by batteries. In such designs, the weight of the batteries can be distributed to match that of a combustion engine to enable the electric or hybrid aircraft to fly similarly to the traditional combustion powered aircraft.

FIG.5Aillustrates a battery module1400usable in an aircraft, such as the aircraft100ofFIGS.1A and1B. The battery module1400can include a lower battery module housing1410, a middle battery module housing1420, an upper battery module housing1430, and a multiple battery cells1440. The multiple battery cells1440can together provide output power for the battery module1400. The lower battery module housing1410, the middle battery module housing1420, or the upper battery module housing1430can include slots, such as slots1422, that are usable to mechanically couple the lower battery module housing1410, the middle battery module housing1420, or the upper battery module housing1430to one another or to another battery module. Supports, such as supports1424(for example, pins or locks), can be placed in the slots to lock the lower battery module housing1410, the middle battery module housing1420, or the upper battery module housing1430to one another or to another battery module.

The battery module1400can be constructed so that the battery module1400is evenly cooled by air. The multiple battery cells1440can include 16 total battery cells where the battery cells are each substantially shaped as a cylinder. The lower battery module housing1410, the middle battery module housing1420, or the upper battery module housing1430can be formed of or include plastic and, when coupled together, have an outer shape substantially shaped as a rectangular prism. The lower battery module housing1410, the middle battery module housing1420, or the upper battery module housing1430can together be designed to prevent a fire in the multiple battery cells1440from spreading outside of the battery module1400.

The battery module1400can have a length of L1, a width of W, and a height of H1. The length of L1, the width of W, or the height of H1can each be 50 mm, 65 mm, 80 mm, 100 mm, 120 mm, 150 mm, 200 mm, 250 mm or within a range defined by two of the foregoing values or another value greater or less than the foregoing values.

FIG.5Billustrates an exploded view of the battery module1400ofFIG.5A. In the exploded view, a plate1450and a circuit board assembly1460of the battery module1400is shown. The plate1450can be copper and may electrically connect the multiple battery cells1440in parallel with one another. The plate1450may also distribute heat evenly across the multiple battery cells1440so that the multiple battery cells1440age at the same rate. The circuit board assembly1460may transfer power from or to the multiple battery cells1440, as well as include one or more sensors for monitoring a voltage or a temperature of one or more battery cells of the multiple battery cells1440. The circuit board assembly1460may or may not provide galvanic isolation to the battery module1400with respect to any components that may be electrically connected to the battery module1400. Each of the multiple battery cells1440can have a height of H2, such as 30 mm, 50 mm, 65 mm, 80 mm, 100 mm, 120 mm, 150 mm or within a range defined by two of the foregoing values or another value greater or less than the foregoing values.

FIG.6Aillustrate a power source1500A formed of multiple battery modules1400ofFIG.14. The multiple battery modules1400of the power source1500A can be mechanically coupled to one another. A first side of one battery module1400can be mechanically coupled to a first side of another battery module1400, and a second side of the one battery module1400that is opposite the first side can be mechanically coupled to a first side of yet another battery module1400. The multiple battery modules1400of the power source1500A can be electrically connected in series with one another. As illustrated inFIG.6A, the power source1500A can include seven of the battery modules1400connected to one another. The power source1500A may, for example, have a maximum power output between 1 KW and 60 KW during operation, a maximum voltage output between 10 V and 120 V during operation, or a maximum current output between 100 A and 500 A during operation.

The power source1500A can include a power source housing1510mechanically coupled to at least one of the battery modules. The power source housing1510can include an end cover1512that covers a side of the power source housing1510. The power source housing1510can have a length of L2, such as 3 mm, 5 mm, 10 mm, 15 mm, 20 mm, 25 mm, 30 mm, 40 mm, 50 mm or within a range defined by two of the foregoing values or another value greater or less than the foregoing values. The width and the height of the power source housing1510can match the length of L1and the width of W of the battery module1400.

The power source1500A can include power source connectors1520. The power source connectors1520can be used to electrically connect the power source1500A to another power source, such as another of the power source1500A.

FIG.6Billustrates a power source1500B that is similar to the power source1500A ofFIG.6Abut with the end cover1512and the upper battery module housings1430of the battery modules1400removed. Because the end cover1512has been removed, a circuit board assembly1514of the power source1500B is now exposed. The circuit board assembly1514can be electrically coupled to the battery modules1400. The circuit board assembly1514can additionally provide galvanic isolation (for instance, 2500 Vrms) for the power source1500B with respect to any components that may be electrically connected to the power source1500B. The inclusion of galvanic isolation in this manner may, for instance, enable grouping of the battery modules1400together so that isolation may be provided to the grouping of the battery modules1400rather than individual modules of the battery modules1400or a subset of the battery modules1400. Such an approach may reduce the costs of construction because isolation can be expensive, and a single isolation may be used for multiple of the battery modules1400.

FIG.7illustrates a group1600of multiple power sources1500A ofFIG.6Aarranged and connected for powering an aircraft, such as the aircraft100ofFIGS.1A and1B. The multiple power sources1500A of the group1600can be mechanically coupled to or stacked on one another. The multiple power sources1500A of the group1600can be electrically connected in series or parallel with one another, such as by a first connector1610or a second connector1620that electrically connects the power source connectors1520of two of the multiple power sources1500A. As illustrated inFIG.7, the group1600can include 10 power sources (for instance, arranged in a 5 row by 2 column configuration). In other examples, a group may include a fewer or greater number of power sources, such as 2, 3, 5, 7, 8, 12, 15, 17, 20, 25, 30, 35, or 40 power sources.

The grouping of the multiple power sources1500A to form the group1600or another different group may allow for flexible configurations of the multiple power sources1500A to satisfy various space or power requirements. Moreover, the grouping of the multiple power sources1500A to form the group1600or another different group may permit relatively easy or inexpensive replacement of one or more of the multiple power sources1500A in the event of a failure or other issue.

FIG.8Aillustrates a perspective view of a nose1700of an aircraft, such as the aircraft100ofFIGS.1A and1B, that includes multiple power sources1710, such as multiple of the power source1500A, for powering a motor1720that operates a propeller1730of the aircraft. The multiple power sources1710can be used to additionally or alternatively power other components of the aircraft. The multiple power sources1710can be sized and arranged to optimize a weight distribution and use of space around the nose1700. The motor1720and the propeller1730can be attached to and supported by a frame of the aircraft by supports, which can be steel tubes, and connected by multiple fasteners, which be bolts with rubber shock absorbers. A firewall1740can provide barrier between the multiple power sources1710and the frame of the aircraft in the event of a first at the multiple power sources1710. An enclosure composed of glass fiber, metal, or mineral composite can be around the multiple power sources1710to protect from water, coolant, or fire.

FIG.8Billustrates a side view of the nose1700ofFIG.8A.

FIG.9Aillustrates a top view of a wing1800of an aircraft that includes multiple power sources1810, such as multiple of the power source1500A, for powering one or more components of the aircraft. The multiple power sources1810can be sized and arranged to optimize a weight distribution and use of space around the wing1800. For example, the multiple power sources1810can be positioned within, between, or around horizontal support beams1820or vertical support beams1830of the wing1800. A relay1840can further be positioned in the wing1800as illustrated and housed in a sealed enclosure. The relay1840may open if there is not a threshold voltage on a breaker panel or if a pilot opens breakers to shut down the multiple power sources1810.

FIG.9Billustrates a perspective view of the wing1800ofFIG.9A.

The battery packs in different portions of the airplanes, for example the battery packs in the wings and/or in the nose, may be connected serially and/or in parallel.

Multi-Coil Motor Control

An electric or hybrid aircraft can be powered by a multi-coil motor, such as an electric motor, in which different coils of the motor power different phases of a modulation cycle for the motor.

As can be seen fromFIG.10, a motor1910can include four different field coils (sometimes also referred to as coils) for generating a torque on a rotor of the motor1910. The different field coils can include a first field coil1902, a second field coil1904, a third field coil1906, and a fourth field coil1908. Each of the different field coils can be independently powered by one or more controllers. The first field coil1902, the second field coil1904, the third field coil1906, and the fourth field coil1908can be respectively powered by a first controller1912, a second controller1914, a third controller1916, and a fourth controller1918. One or more of the first controller1912, the second controller1914, the third controller1916, and the fourth controller1918may be the same controller.

Electrical Power Supply and Electrical Supply Network

FIG.11illustrates an electrical power supply system. The power supply system comprises an electrical power source1, a power converter2and a load3. The electrical power source1provides a DC voltage at its output terminals DC+, DC− with a first potential difference VH.

The level of the DC voltage which is provided by the electrical power source1may account in this embodiment to more than 800V, for example more than 1200 V, such as 1600 V. The electrical power source1can be connected to one or more motor controllers for powering an electrical motor of an aircraft. The one or more motor controllers can face at its input end a voltage that corresponds to the first potential difference VHprovided by the electrical power source1.

The electrical power source1can include a third output N that provides a potential between the potential of the first output terminal DC+ and the second output terminal DC−.

The potential of the third output N can be symmetrical with reference to the potential of the first DC+ or second DC−. In this embodiment, the first potential difference can be divided into two halves around the potential of the third output N as reference.

The voltage between the first output terminal DC+ and the third output terminal N or between the third output terminal N and the second output terminal DC− may account in this embodiment to less than 800V, for example to less than 500 V.

The power converter2can include input terminals and output terminals (not shown).

The power converter2can be connected with its input terminals to the first DC potential output DC+ and to the second DC potential output DC− of the electrical power source1. This means that the power converter2may face at its input terminals the first potential difference VH. In other words, the power converter2may face at its input terminals the full system voltage that is supplied by the electrical power source1.

Alternatively, the power converter2can be connected with its input terminals to the first DC potential output DC+ and to the third DC potential output N or to the third DC potential output N and to the second DC potential output DC− of the electrical power source1. This means that the power converter2may face at its input terminals one half of the first potential difference VH, when the potential of the first DC potential output DC+ and the potential of the second DC potential output DC− is symmetrical with reference to the potential of the third DC potential output N.

The power converter2can feature in addition two output terminals, in particular a first potential output V+ and a second potential output V−.

The power converter2may be configured in this example as a galvanically isolated step-down DC-DC converter. The DC-DC converter can convert the DC voltage with a first potential difference VHfrom a high potential difference at its input terminals to a DC voltage with a second potential difference VLwith a lower potential difference at its output terminals.

In one example, the voltage at the input terminals can account to a DC voltage of more than 800V, for example 1200 V and a DC voltage of 48 V or less, for example 14 V at the output terminal.

In another example, the related voltage can account to a DC voltage of more than 1200 V, for example 1600 V and a DC voltage of 120 V or less, for example 48 V at the output terminal.

In a third example, the voltage at the input terminals can account to a DC voltage of less than 800V, for example 500 V and a DC voltage of 48 V or less, for example 14 V at the output terminal, where the input terminals of the power converter2are connected to the first DC potential output DC+ and to the third DC potential output N of the electrical power source1being at 0 V.

The DC-DC converter can ensure at the same time that a galvanic isolation or at least electrical separation between the electrical circuit on the high voltage side and the load on the low voltage side is maintained. Galvanically isolated electrical systems can prevent, for example, stray currents flowing in the electrical system, causing additional losses. On the other hand, a galvanic isolation can protect people from an isolation failure in the high voltage system, when being in contact with the voltage system. Two isolation failures thus may be required to put people at risk.

The load3can be connected to the output terminals of the power converter2. The load3may absorb the electrical energy that is supplied by the power converter2. The load3can be in this embodiment avionic instruments, digital equipment, electronic equipment, cockpit low voltage battery, and/or other devices that can operate at a lower voltage. Consequently, the load3can be a standard electric consumer, without the ability to supply power to the output terminals of the power converter2. The load3in this specific example may not be a high voltage motor controller.

As indicated before, the electrical power source1can include a third output terminal N. This terminal can be connected to one of the output terminals of the power converter2via a further impedance (not shown). In the illustrated embodiment, the DC output terminal N can be connected to the second potential output V− of the power converter2.

In other words, the third DC output terminal N of the electrical power source1can be connected to the terminal of the power converter2to which the voltage or potential on the output end is referenced to.

The connection between the third DC potential output N and the second potential output V− can include an impedance4. This impedance may be a resistor. The resistance of the resistor may account to more than 100 kΩ, such as 1 MΩ in this example, or more than 1 MΩ.

The third DC potential output N and the second potential output V− can be interconnected without using the impedance4in some embodiments. Arranging the connection without using the impedance4may however lead in case of faults to high currents flowing in this connection or through a human body, which might not be acceptable in the view of electrical safety.

Further devices31,32,33connectable to the electrical system of the aircraft are shown inFIG.11, relating to a further example of this embodiment. An avionic battery31can be connected to the output terminals of the power converter2. The avionic battery31can be charged with electrical energy supplied at the low voltage side of the power converter2. The avionic battery31can also supply the load3in case the energy normally supplied by the power converter2may not be available. The avionic battery31can provide a DC voltage.

A motor controller32can be connected with its input terminals to the first DC potential output DC+ and the second DC potential output DC− of the electrical power source1. The motor controller32can convert a DC voltage from an input end into a three-phase AC voltage with variable frequency and amplitude at an output end to supply a motor33for propelling the aircraft.

The motor controller32and the motor33can include an enclosure or a housing made of an electrically conductive material. The enclosure or housing can be connected to the electrical ground G of the aircraft using a direct connection in the form of a cable connection or a connection including an impedance, such as the impedance4as illustrated inFIG.11.

This is advantageous as the requirements for separating or insulating the high voltage parts, such as the semiconductors of the motor controller32or the windings of the motor33, from the potential of the enclosure can be relaxed for the same reason as set out before. In particular, the first potential difference VHcan be referenced to the second potential output V− of the power converter2. The motor controller32may be supplied at its input end with a DC voltage of 1200 V or higher, such as 1600 V.

FIG.12Aillustrates in a further embodiment a battery pack10including multiple battery modules11,12, Z-1, Z.

The battery pack10can provide a DC voltage to the power converter2. In other words, the battery pack10may act as an electrical power source1that supplies the power converter2with electrical energy.

The DC voltage may be above 450 V, 500 V, 550 V, 650 V. The DC voltage output of the battery pack10may be more than 1200 V, for example, 1600 V.

The DC voltage can be the output between a first DC potential output DC+ and a second DC potential output DC− of the battery pack10. The potential difference between the first DC potential output DC+ and a second DC potential output DC− may provide the DC voltage output from the battery pack10.

The first DC potential output DC+ may provide a first polarity like plus or positive and the second DC potential output DC− may provide a second polarity being different from the first polarity like minus or negative, or vice versa. The DC potential output at the first DC potential output DC+ and the DC potential output at the second DC potential output DC− may have the same absolute value (with different polarities).

The battery pack10can include a plurality of battery modules11,12, Z-1, Z which are connected in parallel or in series in order to obtain the characteristics of the battery pack10regarding to voltage, current, power or capacity.

The battery modules11,12, Z-1, Z can include lithium-based battery cells, e.g. lithium-ion battery cells or lithium-ion polymer batteries. Some of the battery cells may be based on other technologies, e.g. on supercapacitors or others.

According to one embodiment, the battery pack10can include a third DC potential output N.

The third DC potential output N can be at a potential between DC− and DC+, at a potential of 0 V (zero Volt). The third DC potential output N can be at a potential corresponding to the lower one of the potentials of the first and second DC potential output DC− plus the half of the potential difference between the first DC potential output DC+ and second DC potential DC−.

The battery pack10can include a first battery module11, a second battery module12, a second last battery module Z-1and a last battery module Z as shown in theFIG.12A.

The series connection of the first battery module11and the second battery module12may provide a first DC sub-voltage between the first DC potential output DC+ and the third DC potential output N. The series connection of the second last battery module Z-1and the last battery module Z can provide a second DC sub-voltage between the second DC potential output DC− and the third DC potential output N.

The first and second battery modules11,12and the second last battery module Z-1and the last battery module Z can be connected in series between the first DC potential output DC+ and the second DC potential output DC− to provide the DC voltage consisting of the sum of the first DC sub-voltage and the second DC sub-voltage.

The first and second battery module11,12and the second last and the last battery module Z-1, Z each may have a first output terminal and a second output terminal.

The first output terminal of the first battery module11can be connected with the first DC potential output DC+ and the second output terminal of the second battery module12can be connected with the third DC potential output N.

The first output terminal of the second last battery module Z-1can be connected with the third DC potential output N or the second output terminal of the second battery module12and the second output terminal of the last battery module Z can be connected with the second DC potential output DC−.

The battery modules11,12, Z-1, Z can be arranged in a common housing, or in two distinct housings. The connections between the battery modules11,12, Z-1, Z can be an internal connection within a common housing of the battery pack10.

FIG.12Billustrates in a further embodiment of multiple battery strings S1, S2, Sz for providing the electrical power source1. Each of the battery strings S1, S2can include a battery pack10, including multiple battery modules11,12, Z-1, Z as illustrated inFIG.12A or12B.

Each battery string S1, S2can have two terminals for providing a first and a second sub-DC potential output DCs+, DCs−. The first sub-DC potential outputs DCs+ of the battery strings S1, S2may be connected to a positive busbar for providing the first DC potential output DC+. The second sub-DC potential outputs DCs− may be connected to a negative busbar for providing the second DC potential output DC−. The battery strings S1, S2can be connected in parallel through their terminals (sub-DC potential outputs DCs+, DCs−) to provide higher current capabilities.

It can be noticed in the corresponding figure that each battery pack10can be configured with a series connection of battery modules11-Z for providing a high DC voltage output.

Each battery pack10of each battery string S1, S2can provide a DC voltage measured between the first and the second sub-DC potential output DCs+, DCs− of more than 450 V, such as more than 500 V, 550 V, or 650 V.

The DC voltage output of each battery pack10can be more than 1200 V, such as 1600 V.

The voltage between the first and the second sub-DC potential output DCs+, DCs− can correspond to the voltage between the first and second DC potential output DC+, DC−.

Further battery strings Sz can be connected in parallel to the two battery strings S1, S2illustrated. The additional battery strings Sz may be similarly configured as the battery strings S1, S2.

The first and the second DC potential outputs DC+, DC− may supply the power converter2with electrical energy, as illustrated inFIG.11.

The battery pack10of each of the strings S1, S2can include additional electrical components and circuitry. A first switch54, such as a high power contactor, can connect a positive terminal of the battery module11;12to the first sub-DC potential output DCs+. A parallel connected pre-charge circuit, including a second switch55, such as a relay, and a series connected resistor43can limit the inrush current before the first switch54is closed to electrically connect the positive terminal to the first sub-DC potential output DCs+.

Alternatively or in addition, the pre-charge circuit43,55can be connected in parallel to a third switch56, such as a high power contactor, providing similar or the same switching functionality as the first switch54.

The resistor43can have a low to medium ohmic value, such as between 100Ω and 10 kΩ. The resistance of the resistor43can be more than an order of magnitude lower than the resistance of the impedance4.

The third switch56can connect a negative terminal of the battery module11;12to the second sub-DC potential output DCs−. For pre-charging, the second and the third switches55,56can be closed synchronously. For disconnecting the string S1, S2, the first and the third switches54,56can be opened simultaneously.

The switches54-56can be connected to and controlled by a controller. The controller can be configured as illustrated inFIG.16and described in more detail in the present disclosure. Alternatively, the switches54-56can be connected to and controlled by a higher level controller of the aircraft100.

The battery pack10ofFIG.12Aor the battery packs15,16ofFIG.13can also be configurated with the switches and the pre-charging circuit, as set out hereinbefore.

Each battery pack10of one string S1, S2can be arranged with a terminal to provide a third DC potential output N. The third DC potential output terminal can be connected through an impedance4to the electrical ground G of the aircraft100. One or more or all of the battery string S1, S2can have with an impedance4.

Each impedance4can be or include a resistor with a rating, for example, of 1 MΩ or higher.

FIG.13illustrates a series connection of two battery packs15,16. Each of the battery packs15,16can be arranged for example as disclosed in the embodiments ofFIGS.12A and12B.

The battery packs15,16are arranged and can provide an intermediate DC voltage between a first DC potential output DC+ and a second DC potential output DC− for supplying the power converter2with electrical energy. In other words, the battery packs15,16can be arranged to form a power source, such as the electrical power source1that supplies the power converter2with electrical energy.

The DC voltage between said DC potential outputs DC+, DC− can be above 450 V, 500 V, 550 V, or 650 V. The DC voltage can account to more than 1200 V, such as 1600 V.

The two battery packs15,16can be connected in series. Each of the two battery packs may be provided with two terminals (not shown). The first terminal can be at a first positive potential, whereas the second terminal can be at a negative potential with reference to the potential of the first terminal.

The first terminal of the second battery pack16may provide a first sub-DC potential output DCs+, whereas the second terminal of the first battery pack15may provide a second sub-DC potential output DCs−. The two sub-DC potential outputs DCs+, DCs− can be interconnected to provide a third DC potential output N at the interconnection.

The first terminal of the first battery pack15may provide the first DC potential output DC+ and the second terminal of the second battery pack16may provide the second DC potential output DC−. The power converter2can be connected to the said DC potential outputs DC+, DC−.

The two battery packs15,16can be arranged in one common housing.

Alternatively, a first battery pack15can be arranged or placed in one location of the aircraft100, whereas a second battery pack16can be placed in another location of the aircraft100. The first location can be a wing portion of the first wing, whereas a second location can be a wing portion of the second wing of the aircraft100. Both battery packs may also be located in the same wing.

In this embodiment, the first sub-DC potential output DCs+ can be connected to the second sub-DC potential output DCs−, using a cable that is laid through the aircraft100, as the two battery packs15,16may be physically separated from each other. This cable can be configured with a tab, so that the second potential output V− of the power converter2can be connected to said third DC potential out N.

FIG.14illustrates the internal structure of the power converter2with its electrical isolation.

The power converter2may be a DC-DC converter that can be connected with its input terminals to the first DC and second DC potential output DC+, DC− of the electrical power source1. The electrical power source1can be arranged either in the form of a set of battery modules or as a set of battery packs, as explained before. The power converter2can be further connected at its output terminals to the load3. It is also possible to use two converters, one connected to the second DC potential output DC− and to the third DC potential out N and the other one connected to the third DC potential out N and to the first DC potential output DC+ while the outputs terminal can be put together to form an output terminal of a second potential output V− that is connected to the third DC potential output of the electrical power source1via a impedance.

The power converter2can convert a voltage with a first potential difference VHat its input terminals into a voltage with a second potential difference VLat its output terminals. The voltage at the input terminals can be greater than the voltage at the output terminals of the power converter2.

The output terminal, in particular, the second potential output of the power converter2is connected to the electrical ground G of the aircraft100. The third DC potential output of the electrical power source can be connected to the second potential output V− or electrical ground G of the aircraft via an impedance (not shown).

The power converter2can include a high voltage primary side21and a low voltage secondary side23. The high voltage primary side may be electrical isolated from the low voltage secondary side using an insulating barrier25. The insulating barrier25can be a physical component configured to isolate the primary side from the secondary side and vice versa.

The insulating barrier25can be a transformer that separates the input end of the DC-DC converter from the output end. The electrical supply network may operate in a fault state, when the insulating barrier25may not be separating the high voltage primary side21from the low voltage secondary side23.

FIG.15illustrates a power converter2with its enclosure22and a cable26connecting the power converter2to one of the DC potential DC+, DC− outputs of the electrical power source1.

The power converter2, which can be a DC-DC converter, can be placed in an enclosure22. This enclosure22can be constructed of an electrically conductive material, such as steel or aluminium. The enclosure22may protect the components arranged inside the enclosure22against water or dust, as the enclosure22may feature an ingress protection. The enclosure22can connected to electrical ground G of the aircraft100and may be connected to the third DC potential output of the electrical power source1via an impedance (not shown).

The electrical cable26can connect one of the electrical outputs of the electrical power source1to the high voltage primary side21of the power converter2. The cable26can enter the enclosure22via a feed-through24. The electrical cable26may include a wire insulation (not shown) to electrical isolate the cable26from other potentials, such as the potential of the enclosure22.

The enclosure22can be made of electrically conductive material, and the enclosure22may be connected to electrical ground G or to the third DC potential output N of the electrical power source1. The insulation of the cable26can be configured to withstand one half of the potential difference of the first potential difference VH. Same for the motor coils and the motor housing.

The cable26may get into contact with the electrically conductive enclosure22, in case the insulation of the cable26is broken. This can lead to a fault in the electrical supply system or supply network that can be detected using the method as disclosed herein after.

FIG.16illustrates an electrical supply network. The electrical power supply network can include an electrical power source1, a power converter2and a load3that is connected to the output terminals of said power converter2. The power converter2is connected to the electrical power source1. The electrical power supply network can include in addition a first connection for connecting the third DC potential output N of the power source1to the second potential output of the power converter2.

The electrical power supply network furthermore can include a second and a third connection for connecting the first DC potential output DC+ and the second DC potential output DC− of the electrical power source1to the second potential output of the power converter2. The second potential output of the power converter2can be connected to the electrical ground G of the aircraft100.

Each of the connections may include an impedance4,41,42for limiting a current flow. Each of the impedances4,41,42can include a resistor with a rating for example of 1 MΩ or more. The second and third connection can furthermore include a switch52,53configured to interrupt or establish the connection.

The power converter2can include a contactor51at its input end, configured to separate the power converter2from the electrical power source1.

The electrical power supply network further can include a controller6, for example a microcontroller. The controller6may have analogue inputs, configured to convert a voltage level into a digital value. The controller6can include analogue outputs configured to convert a digital value into voltage level.

The electrical power supply network further may include three voltage sensors (not shown) that are connected to each of the three impedances4,41,42and configured to measure a voltage across each of the said impedances4,41,42. The analogue inputs of the controller6can be connected to the voltage sensors. Two of the analogue outputs of the controller6can be connected to the switches52,53. One further analogue output can be connected to the contactor51.

The controller6can measure a voltage across the impedances4,41,42and control the contactor51or the switches52,53between an open position (which may relate to a non-conductive state) and a closed position (which may relate to a conductive state).

The controller6can execute a software or perform a control method that implements the following steps, such as in the order as disclosed hereinafter.

A control method for testing the insulation integrity of the electrical power supply system or the electrical supply network can be executed by the controller6during preparation of the aircraft100for takeoff.

Alternatively or in addition, the insulation integrity can be tested by executing the method as outlined hereinafter when the aircraft100is cruising.

In an initial state of the electrical power supply network the power converter2may be connected to the electrical power source1as the contactor is in a closed position.

The first switch52and the second switch53can be in an open position, meaning that the connection between one of the DC potential outputs DC+, DC− of the electrical power source and the second potential output V− of the power converter2is interrupted. The electrical power source1may supply the power converter2with electrical energy, whereas the power converter2may convert the electrical energy from the high voltage primary side21into electrical energy on the low voltage secondary side23.

For testing the insulation integrity of the electrical supply network, the controller6may determine the voltage across the first impedance4by measuring the voltage at the analogue input that is connected to the related voltage sensor.

Under normal operational conditions the voltage across the first impedance4may be zero or close to zero, as the electrical supply network on the side of the electrical power source1is electrically insulated from the electrical supply network on the secondary side of the power converter2.

In a further, potentially subsequent step for testing the insulation integrity of the electrical supply network, the controller6may control the first switch52from an open position to a closed position, to arrange the connection between the first DC potential output DC+ of the electrical power source1and the second potential output V− of the power converter2.

The controller6may determine the voltage across the first impedance4by capturing the voltage at the analogue input that is connected to the related voltage sensor.

The controller6may further determine the voltage across the second impedance41by capturing the voltage at the analogue input that is connected to the related voltage sensor.

The controller6can, in addition in a subsequent step, control the first switch52from a closed position to an open position to interrupt the connection between the first DC potential output DC+ of the electrical power source1and the second potential output V− of the power converter2.

Under normal operational conditions, the voltage across the first impedance4may be equal to the voltage across the second impedance41, as the electrical supply network on the side of the electrical power source1is electrically insulated from the electrical supply network on the secondary side of the power converter2.

Furthermore, a comparison of the two voltage measurements may indicate that the electrical supply network on the side of the electrical power source1does not have an insulation fault.

In a further, potentially subsequent step for testing the insulation integrity of the electrical supply network, the controller6may control the second switch53from an open position to a closed position, to arrange the connection between the second DC potential output DC− of the electrical power source1and the second potential output V− of the power converter2.

The controller6may determine the voltage across the first impedance4by capturing the voltage at the analogue input that is connected to the related voltage sensor.

The controller6may further determine the voltage across the third impedance42by capturing the voltage at the analogue input that is connected to the related voltage sensor.

In a subsequent step, the controller6can control the second switch53from a closed position to an open position to interrupt the connection between the second DC potential output DC− of the electrical power source1and the second potential output V− of the power converter2.

Under normal operational conditions, the voltage across the first impedance4may be equal to the voltage across the third impedance42, as the electrical supply network on the side of the electrical power source1is electrically insulated from the electrical supply network on the secondary side of the power converter2.

Furthermore, a comparison of the two voltage measurements may indicate that the electrical supply network on the side of the electrical power source1does not have an insulation fault.

A comparison of the two measurements of the voltage across the second impedance41and the third impedance42may indicate that there is no insulation fault apparent when the two voltage measurements are the same values.

When the impedances are configured to have the same electrical properties, the potential of the first DC potential output DC+ can be symmetrical to the potential of the second DC potential output DC− the voltage across the two said impedances (second impedance41and third impedance42) and thus be at the same level.

In case the impedances are configured with different electrical properties, such as a different resistance, the introduction of a threshold value or reference value to which the measurements are compared, the measurements may not be directly comparable to each due to deviating electrical properties.

Comparing the voltage across the three impedances4,41,42with each other provides the advantage that insulation faults between the high voltage primary side21and the low voltage secondary side23can be detected in addition to insulation faults on the side of the electrical power source1.

In particular, insulation faults in the electrical power source1can be detected. Such insulation faults can be caused, for example, by damaged battery modules11,12, Z-1, Z placed in a housing of the electrical power source1or broken cable insulation.

The controller6can communicate with other control devices in the aircraft100via a communication bus or a hard wired connection.

In case one of the measurements of the voltage deviates from a measurement that is expected during normal operation, the controller6may alert other control devices in the aircraft, that the insulation integrity of the electrical power supply network or the electrical power source is violated.

The controller6may initiate further actions, such as to open the contactor51, for disconnecting the power converter2from the electrical power source1in case said insulation integrity violation is detected.

Additional Features and Terminology

Although examples provided herein may be described in the context of an aircraft, such as an electric or hybrid aircraft, one or more features may further apply to other types of vehicles usable to transport passengers or goods. For example, the one or more futures can be used to enhance construction or operation of automobiles, trucks, boats, submarines, spacecrafts, hovercrafts, or the like.

Many other variations than those described herein will be apparent from this disclosure. For example, depending on the embodiment, certain acts, events, or functions of any of the algorithms described herein can be performed in a different sequence, can be added, merged, or left out altogether (for example, not all described acts or events are necessary for the practice of the algorithms). Moreover, in certain embodiments, acts or events can be performed concurrently, for instance, through multi-threaded processing, interrupt processing, or multiple processors or processor cores or on other parallel architectures, rather than sequentially. In addition, different tasks or processes can be performed by different machines or computing systems that can function together.

Unless otherwise specified, the various illustrative logical blocks, modules, and algorithm steps described herein can be implemented as electronic hardware, computer software, or combinations of both. To clearly illustrate this interchangeability of hardware and software, various illustrative components, blocks, modules, and steps have been described above generally in terms of their functionality. Whether such functionality is implemented as hardware or software depends upon the particular application and design constraints imposed on the overall system. The described functionality can be implemented in varying ways for each particular application, but such implementation decisions should not be interpreted as causing a departure from the scope of the disclosure.

Unless otherwise specified, the various illustrative logical blocks and modules described in connection with the embodiments disclosed herein can be implemented or performed by a machine, a microprocessor, a state machine, a digital signal processor (DSP), an application specific integrated circuit (ASIC), a FPGA, or other programmable logic device, discrete gate or transistor logic, discrete hardware components, or any combination thereof designed to perform the functions described herein. A hardware processor can include electrical circuitry or digital logic circuitry configured to process computer-executable instructions. In another embodiment, a processor includes an FPGA or other programmable device that performs logic operations without processing computer-executable instructions. A processor can also be implemented as a combination of computing devices, e.g., a combination of a DSP and a microprocessor, a plurality of microprocessors, one or more microprocessors in conjunction with a DSP core, or any other such configuration. A computing environment can include any type of computer system, including, but not limited to, a computer system based on a microprocessor, a mainframe computer, a digital signal processor, a portable computing device, a device controller, or a computational engine within an appliance, to name a few.

Unless otherwise specified, the steps of a method, process, or algorithm described in connection with the embodiments disclosed herein can be embodied directly in hardware, in a software module stored in one or more memory devices and executed by one or more processors, or in a combination of the two. A software module can reside in RAM memory, flash memory, ROM memory, EPROM memory, EEPROM memory, registers, hard disk, a removable disk, a CD-ROM, or any other form of non-transitory computer-readable storage medium, media, or physical computer storage known in the art. An example storage medium can be coupled to the processor such that the processor can read information from, and write information to, the storage medium. In the alternative, the storage medium can be integral to the processor. The storage medium can be volatile or nonvolatile. The processor and the storage medium can reside in an ASIC.

Conditional language used herein, such as, among others, “can,” “might,” “may,” “e.g.,” and the like, unless specifically stated otherwise, or otherwise understood within the context as used, is generally intended to convey that certain embodiments include, while other embodiments do not include, certain features, elements or states. Thus, such conditional language is not generally intended to imply that features, elements or states are in any way required for one or more embodiments or that one or more embodiments necessarily include logic for deciding, with or without author input or prompting, whether these features, elements or states are included or are to be performed in any particular embodiment. The terms “comprising,” “including,” “having,” and the like are synonymous and are used inclusively, in an open-ended fashion, and do not exclude additional elements, features, acts, operations, and so forth. Also, the term “or” is used in its inclusive sense (and not in its exclusive sense) so that when used, for example, to connect a list of elements, the term “or” means one, some, or all of the elements in the list. Further, the term “each,” as used herein, in addition to having its ordinary meaning, can mean any subset of a set of elements to which the term “each” is applied.