Power distribution system for an aircraft

An aircraft power distribution system includes a first DC power distribution bus, a second DC power distribution bus, and a DC power source coupled with at least one of the first or second DC power distribution buses, wherein the DC power distribution buses are electrically coupled by a plurality of electrical couplings.

BACKGROUND OF THE INVENTION

Power systems, especially power systems in aircraft, manage the supplying of power from power sources, such as generators, to electrical loads. In aircraft, gas turbine engines are used for propulsion of the aircraft, and typically provide mechanical power which ultimately powers a number of different accessories such as generators, starter/generators, permanent magnet alternators (PMA), fuel pumps, and hydraulic pumps, e.g., equipment for functions needed on an aircraft other than propulsion. For example, contemporary aircraft need electrical power for avionics, motors, and other electric equipment. A generator coupled with a gas turbine engine will convert the mechanical power of the engine into electrical energy which is distributed throughout the aircraft by electrically coupled nodes of the power distribution system. The power distribution system may fail at any of the coupled nodes, which may interrupt the electrical power distribution, as well as any equipment reliant on that power.

BRIEF DESCRIPTION OF THE INVENTION

In one aspect, an aircraft power distribution system includes a first DC power distribution bus and a second DC power distribution bus, a DC power source coupled with at least one of the first or second DC power distribution buses, a plurality of electrical couplings between the first and second DC power distribution buses, and a solid state power controller positioned in-line on each electrical coupling. Each solid state power controller further includes two power switches in a back-to-back configuration, and each power switch comprising a field-effect transistor (FET) connected across a Schottky diode. The solid state power controllers selectively couple and decouple the at least first and second DC power distribution buses across the plurality of electrical couplings.

In another aspect, a method of operating an aircraft power distribution system having at least one DC power source coupled with a first DC power distribution bus, and a second DC power distribution bus, and a solid state power controller positioned on each of a plurality of electrical couplings between the at least first and second DC power distribution buses, the method includes determining if a fault occurs in at least one electrical couplings between the first DC power distribution bus and the second DC power distribution bus, to define a faulted connection, and controlling the solid state power controllers, based on the determination of a fault, to selectively decouple the faulted connection between the first DC power distribution bus and the second DC power distribution bus, and to selectively recouple the first DC power distribution bus with the second DC power distribution bus via another electrical coupling other than the one with the faulted connection. The time to selectively recouple the first DC power distribution bus with the second DC power distribution bus is less than the time for an electrical load, coupled with the second DC distribution bus, to enter into a power interruption reset mode.

DESCRIPTION OF EMBODIMENTS OF THE INVENTION

The described embodiments of the present invention are directed to an electrical power distribution system for an aircraft, which enables production and distribution of electrical power from a turbine engine, preferably a gas turbine engine, to the electrical loads of the aircraft.

As illustrated inFIG. 1, an aircraft10is shown having at least one gas turbine engine, shown as a left engine system12and a right engine system14. Alternatively, the power system may have fewer or additional engine systems. The left and right engine systems12,14may be substantially identical, and are shown further comprising at least one electric machine, such as a generator18. The aircraft is shown further comprising a plurality of power-consuming components, or electrical loads20, for instance, an actuator load, flight critical loads, and non-flight critical loads. Each of the electrical loads20are electrically coupled with at least one of the generators18.

In the aircraft10, the operating left and right engine systems12,14provide mechanical energy which may be extracted via a spool, to provide a driving force for the generator18. The generator18, in turn, provides the generated power to the electrical loads20for load operations. Additional power sources for providing power to the electrical loads20, such as emergency power sources, ram air turbine systems, starter/generators, or batteries, are envisioned. It will be understood that while one embodiment of the invention is shown in an aircraft environment, the invention is not so limited and has general application to electrical power systems in non-aircraft applications, such as other mobile applications and non-mobile industrial, commercial, and residential applications.

FIG. 2illustrates a schematic block diagram of a power distribution system22for an aircraft having multiple engine systems, shown including the left engine system12and the right engine system14, connected by an electrical coupling23. The power distribution22system is shown further including a system controller24, one or more non-engine power sources, shown as an auxiliary power unit (APU)26having an auxiliary power contactor (APC)28and an external ground power source30having an external power contactor (EPC)32, and a tie bus33electrically connecting the left engine system12, right engine system14, APU26, and external ground power source30, in parallel. Each of the APC28and EPC32are configured to selectively couple the respective APU26and external ground power source30to the tie bus33. Additional power sources may be envisioned in addition to, or replacing one or more of the APU26and/or external ground power source30. For instance, an emergency battery system, normal operation battery or battery bank system, fuel cell system, and/or ram air turbine system may be included in the power distribution system22, wherein each may be electrically coupled with the tie bus33, in a parallel configuration.

The left engine system12is shown comprising a first DC power distribution bus34, a second DC power distribution bus36, a first integrated converter controller (ICC)38, a second ICC40, a first generator42capable of generating AC power, and a second generator44capable of generating AC power. The first DC power distribution bus34is connected, via electrical couplings, with at least one electrical load20, the tie bus33, the second DC power distribution bus36, and the first ICC38, which is further electrically coupled with the first generator42. The second DC power distribution bus34is connected, via electrical couplings, with at least one electrical load20and the second ICC40, which is further electrically coupled with the second generator44. Each ICC38,40may additionally provide a fault indication if an error occurs in the ICC38,40, or if the ICC38,40operates outside of operational expectations. Each DC power distribution bus34,36may be configured to provide, for instance 28 VDC or 270 VDC.

The left engine system12may further comprise a first solid state power controller (SSPC)46positioned in-line on the electrical coupling connecting the first DC power distribution bus34with the tie bus33, such that the first SSPC46is between the bus34and the non-engine power sources26,30, and a second SSPC48positioned in-line on the electrical coupling connecting the first DC power distribution bus34with the second DC power distribution bus36.

The left and right engine systems12,14may be substantially identical. Thus, the right engine system14is shown comprising a third DC power distribution bus50, a fourth DC power distribution bus52, a third integrated converter controller (ICC)54, a fourth ICC56, a third generator58capable of generating AC power, and a fourth generator60capable of generating AC power. The third DC power distribution bus50is connected, via electrical couplings, with at least one electrical load20and the third ICC54, which is further electrically coupled with the third generator58. The fourth DC power distribution bus52is connected, via electrical couplings, with at least one electrical load20, the tie bus33, the third DC power distribution bus50, and the fourth ICC56, which is further electrically coupled with the fourth generator60. Each ICC54,56may additionally provide a fault indication if an error occurs in the ICC54,56, or if the ICC54,56operates outside of operational expectations. Each DC power distribution bus50,52may be configured to provide, for instance 28 VDC or 270 VDC.

The right engine system14may further comprise a third SSPC62positioned in-line on the electrical coupling connecting the fourth DC power distribution bus52with the tie bus33, such that the third SSPC62is between the bus34and the non-engine power sources26,30, and a fourth SSPC64positioned in-line on the electrical coupling connecting the third DC power distribution bus50with the fourth DC power distribution bus52. The power distribution system22further comprises a fifth SSPC66positioned in-line on the electrical coupling connecting the second DC power distribution bus36of the left engine system12with the third DC power distribution bus50of the right engine system14. The combined configuration of the tie bus33, the SSPCs46,48,62,64,66, and the DC power distribution buses34,36,50,52defines a ring-type bus configuration74.

Each SSPC46,48,62,64,66comprises two power switches68in a back-to-back configuration, with each power switch68further comprising a field-effect transistor (FET)70(illustrated as a switch) connected across a diode, such as a Schottky diode72. Stated another way, the FET70and Schottky diode72of each power switch68are configured in parallel. The FET70may further comprise a metal-oxide-semiconductor field-effect transistor (MOSFET), such as silicon carbide or gallium nitride MOSTFET, to allow for high power and high speed switching operations. Additionally, it is envisioned each SSPC46,48,62,64,66may be configured with power sensing capabilities to provide a fault indication if a fault occurs within, or on either side of, the SSPC46,48,62,64,66.

As illustrated, the back-to-back configuration is defined by an arrangement of the power switches68such that the Schottky diode72of each switch68is forward-biased away from the opposing switch68. The back-to-back configuration of the power switches68provides each SSPC46,48,62,64,66a selectively energized, or conducting mode, and a selectively de-energized, or non-conducting mode. During the energized mode, the FET70of each power switch68is controlled such that the SSPC46,48,62,64,66allows for electrical coupling between two DC power distribution buses, for instance, the first and second DC power distribution buses34,36. During the de-energized mode, the FET70of each power switch68is controlled such that the SSPC46,48,62,64,66prevents electrical coupling between two DC power distribution buses. Additionally, the location of the first SSPC46and third SSPC62allow these SSPCs46,62to selectively couple and decouple their respective first and fourth DC power distribution buses34,52from the tie bus33, and consequently, the non-engine power sources26,30, during their respective energizing and non-conducting modes.

The system controller24of the power distribution system22is electrically coupled with each of the SSPCs46,48,62,64,66, each ICC38,40,54,56, the APC28, and the EPC32such that the controller24may be in bidirectional communication with, and capable of controlling, each of the aforementioned components. The system controller24may, for instance, independently control each of the aforementioned components or control a plurality of components as a group, as necessary.

While a left engine system12and a right engine system14are shown, alternative embodiments are envisioned having more engine systems for the aircraft. Each engine system may be substantially identical to those illustrated, and may operate in substantially similar fashions. Additionally, while generators42,44,58,60are described, it is envisioned that one or more generators42,44,58,60may alternatively be replaced by a starter/generator, for providing left or right engine system12,14starting functionality. Additionally, alternative embodiments are envisioned wherein each engine system12,14may have more or fewer generators, ICCs, and DC power distribution buses, so long as an SSPC is positioned in-line with each electrical coupling between DC power distribution buses, and in-line with each electrical coupling between a DC power distribution bus and a non-engine power source.

During operation of the power distribution system22, the running gas turbine engines of the left and right engine systems12,14provide mechanical power used by each of the respective first and second generators42,44and third and fourth generators54,56to generate an AC power output. The AC power output of each generator is supplied to a respective ICC38,40,54,56, each of which is controlled by the system controller24to act as an AC to DC rectifier, provide a controlled DC power output, such as 270 VDC, to each respective DC power distribution bus34,36,50,52, which is used to power the electrical loads20.

The DC power distribution buses34,36,50,52may additionally supply power to, or receive power from each other through a plurality of selective electrical coupling paths between each DC power distribution buses34,36,50,52, due to the ring-type bus configuration74. Each of the pluralities of electrical coupling paths between DC power distribution buses34,36,50,52may be controlled by the system controller24selectively energizing or de-energizing each individual or plurality of SSPCs46,48,62,64,66, via a control signal, during normal bus switching operation. For example, the first DC power distribution bus34may supply DC power to the second DC power distribution bus36via at least two electrical coupling paths controlled by the selective coupling or decoupling of the system controller24: directly through the second SSPC48; and around the ring-type bus configuration74, via the first SSPC46, tie bus33, third SSPC62, fourth DC power distribution bus52, fourth SSPC64, third DC power distribution bus50, fifth SSPC66, to the second DC power distribution bus36.

In this sense, the system controller24may be capable of controlling the power distribution system22to redirect power distribution. For example, the system controller24may determine if a fault occurs, in at least one DC power distribution bus34,36,50,52, SSPC46,48,62,64,66, ICC38,40,54,56, or generator42,44,58,60, by way of the bidirectional communication between the controller24and the aforementioned components capable of indicating a fault. This determination of a fault may further distinguish between a clearable fault and a permanent fault, such as a short in an electrical coupling. If a fault is determined to have occurred, the system controller24may define the particular faulted component or connection.

After the system controller24determines a fault has or is occurring, it may selectively decouple or isolate the faulted component or connection from the power distribution system22, and, if possible, re-route or recouple the power distribution path through another electrical coupling other than the faulted component.

For example, if an electrical fault occurs, the system controller24may be alerted to a faulted condition via a fault indictor from one or more of the first SSPC46, the second SSPC48, the fifth SSPC66, first ICC38, or second ICC40. The system controller24may then use the fault indicators to determine or verify if a fault is occurring, and where a fault is occurring, if necessary. For example, the system controller24may determine and define a fault is occurring at the second SSPC48, based on the fault indicators received.

The controller24may further determine if the fault is a permanent fault or a clearable fault based on the fault indicators received. If the fault indicators received indicate a permanent failure of the second SSPC48, the system controller24may selectively control the SSPCs46,48,62,64,66, to decouple the second SSPC48from the first and second DC power distribution buses34,36, and couple the first, third, fourth, and fifth SSPCs46,62,64,66to provide an alternate power distribution path between the buses34,36. In this example, the power distribution system22may selectively decouple (via the second SSPC48) and recouple (via SSPCs46,62,64,66) the first and second DC power distribution buses34,36in less than the time for an electrical load20to detect a potential power interruption, and thus, prevent the electrical load20from entering into a power interruption reset mode. One non-limiting example of the time it may take to collectively decouple and recouple the first and second DC power distribution buses34,36, via another electrical path, may be less than 50 milliseconds.

In an alternate operation of the power distribution system22, wherein the fault indicators received by the system controller24indicate a clearable fault of, for example, the second SSPC48, the system controller24may selectively control the second SSPC48to decouple the first and second DC power distribution buses34,36, and then selectively control the second SSPC48to recouple the buses34,36such that the decoupling and recoupling resets or clears the fault indication. Again, it is envisioned that the decoupling and recoupling of the first and second DC power distribution buses34,36via the second SSPC48occurs in less than the time for an electrical load20to detect a potential power interruption, and thus, prevent the electrical load from entering into a power interruption reset mode. One non-limiting example of the time it may take to collectively decouple and recouple the first and second DC power distribution buses34,36may be less than 50 milliseconds.

Additionally during operation of the power distribution system22, the non-engine power sources26,30may provide primary or supplement power to one or more DC power distribution buses34,36,50,52, via the tie bus33and the first SSPC46and/or third SSPC62. For instance, the system controller24may control the APC28to electrically couple the APU26with the tie bus33to supply supplemental power to the power distribution system22during transient moments of high power requirements. In another instance, the system controller24may control the EPC32to electrically couple the external ground power source30to the tie bus33to supply starting power to the tie bus33, and consequently to a starter/generator, to provide starting functionality for the left or right engine system12,14.

In this sense, the system controller24may additionally be capable of controlling the power distribution system22coupled with a non-engine power source26,30in the event a fault occurs. Similar to the examples above, if either the first or fourth DC power distribution bus34,52fails due to a fault, the system controller24may controllably decouple the bus34,52from the power distribution system22by controlling the corresponding first and second SSPCs46,48, or third and fourth SSPCs62,64in order to isolate the faulted bus34,52from the power distribution system22while still allowing the non-engine power sources26,30to supply power to the remaining, non-faulted buses. Similarly, in an example wherein the third DC power distribution bus50generates a permanent or clearable fault while a non-engine power source26,30is supplying power, the system controller24may isolate the bus50by controlling the fourth and fifth SSPCs64,66to decouple the bus33from the power distribution system22.

Also similar to the method described above, it is envisioned that the power distribution system22may determine if a DC power distribution bus should be isolated from the tie bus33or the system22due to a fault, then control the SSPCs46,48,62,64,66, based on this determination, to selectively decouple the faulted DC power distribution bus from the tie bus33or system22in less than the time for an electrical load20to detect a potential power interruption, and thus, prevent the electrical load20from entering into a power interruption reset mode. Also similar to the method described above, if the power distribution system22determines the DC power distribution bus fault may be cleared, the system controller24may selectively decouple and then recouple the faulted DC power distribution bus to the tie bus33or system22, such that the decoupling/recoupling clears the fault, in less than the time for an electrical load20to detect a potential power interruption, and thus, prevent the electrical load20from entering into a power interruption reset mode.

The embodiments disclosed herein provide a power distribution system. One advantage that may be realized in the above embodiments is that the above described embodiments have superior weight and size advantages over the conventional type power distribution systems due to reduced weight and volume requirements of the solid state power controllers located in bus sharing equipment. Another advantage that may be realized in the above embodiments is that the plurality of selectable power distribution paths provides a robust power distribution system with improved immunity from one or more electrical faults, reducing the likelihood of partial or total aircraft electrical failure. Yet another advantage of the above described embodiments is that the operation of coupling and decoupling the DC power distribution buses by solid state FETs provide for increased reliability because of the lack of mechanical componentry, and thus, reduces the likelihood of mechanical failure in the power distribution system. Even yet another advantage of the above described embodiments is that the embodiments provide a power distribution system with high speed switching that provides detection of faults, and alternate routing or clearing of the said faults, in less time than it takes for the electrical loads to enter into a power interruption reset mode, which provides for uninterrupted electrical load operation despite an electrical fault.

When designing aircraft components, important factors to address are size, weight, and reliability. The above described power distribution system has a decreased number of parts as the system will be able to provide regulated power distribution, making the complete system inherently more reliable. This results in a lower weight, smaller sized, increased performance, and increased reliability system. The lower number of parts and reduced maintenance will lead to a lower product costs and lower operating costs. Reduced weight and size correlate to competitive advantages during flight.

To the extent not already described, the different features and structures of the various embodiments may be used in combination with each other as desired. That one feature may not be illustrated in all of the embodiments is not meant to be construed that it may not be, but is done for brevity of description. Thus, the various features of the different embodiments may be mixed and matched as desired to form new embodiments, whether or not the new embodiments are expressly described. All combinations or permutations of features described herein are covered by this disclosure.