System and method for controlling an aircraft electrical distribution system
An assembly for an aircraft includes an aircraft electrical distribution bus, a battery, and a control system. The battery includes a plurality of battery strings. Each battery string includes a string contactor. The control system is in signal communication with the string contactor of each battery string. The control system is configured to control a position of the string contactor of each battery string in a closed string position or an open string position. The control system is configured to: detect a fault in a faulted first battery string of the plurality of battery strings and position the string contactor of the faulted first battery string in the open string position, in response to the detected fault in the faulted first battery string and maintain the string contactor of at least one unfaulted second battery string of the plurality of battery strings in the closed string position.
1. An assembly for an aircraft, the assembly comprising:
an aircraft electrical distribution bus;
a battery including a plurality of battery strings and a main battery contactor,
the plurality of battery strings electrically connected in parallel, each battery string including a string contactor, the string contactor positionable in a closed string position or an open string position, the string contactor in the closed string position configured to direct electrical power from the respective battery string to the aircraft electrical distribution bus, the string contactor in the open string position configured to electrically isolate the respective battery string from the aircraft electrical distribution bus, and
the main battery contactor electrically connected in series with the string contactor of each battery string, the main battery contactor positionable in a closed battery position or an open battery position, the main battery contactor in the closed battery position configured to direct electrical power from the battery to the aircraft electrical distribution bus, the main battery contactor in the open battery position configured to electrically isolate the battery from the aircraft electrical distribution bus;
a thermal management system for the battery, the thermal management system configured to direct a cooling medium to each battery string of the plurality of battery strings; and
a control system in signal communication with the string contactor of each battery string, the control system configured to control a position of the string contactor of each battery string in the closed string position or the open string position, the control system including a processor in communication with a non-transitory memory storing instructions, which instructions when executed by the processor, cause the processor to:
detect a fault in a faulted first battery string of the plurality of battery strings;
position the string contactor of the faulted first battery string in the open string position, in response to the detected fault in the faulted first battery string, and maintain the string contactor of at least one unfaulted second battery string of the plurality of battery strings in the closed string position; and
control the thermal management system to increase a flow rate of the cooling medium to the at least one unfaulted second battery string in response to detection of the fault in a faulted first battery string.
2. The assembly of claim 1 , wherein control of a position of the main battery contactor in the closed battery position and the open battery position is independent of the control system.
3. The assembly of claim 1 , wherein the instructions, when executed by the processor, further cause the processor to determine an electrical loading limit for the at least one unfaulted second battery string with the string contactor of the faulted first battery string in the open string position.
4. The assembly of claim 3 , wherein the instructions, when executed by the processor, further cause the processor to generate a warning identifying the determined electrical loading limit.
5. The assembly of claim 1 , wherein the instructions, when executed by the processor, further cause the processor to decrease an electrical loading of the aircraft electrical distribution bus in response to detection of the fault in a faulted first battery string.
6. The assembly of claim 1 , wherein each battery string includes a plurality of battery modules electrically connected in series.
7. A method for operating an electrical distribution system for an aircraft, the method comprising:
supplying electrical power to an aircraft electrical distribution bus with a battery, the battery including a plurality of battery strings and a main battery contactor, each battery string including a string contactor electrically connected in series with the main battery contactor;
detecting a fault in at least one faulted first battery string of the plurality of battery strings;
positioning the string contactor of the faulted first battery string in an open string position, in response to the detected fault in the faulted first battery string, and maintaining the string contactor of at least one unfaulted second battery string of the plurality of battery strings in a closed string position; and
determining, with a control system, an electrical loading limit for the at least one unfaulted second battery string with the string contactor of the at least one faulted first battery string in the open string position.
8. The method of claim 7 , wherein the plurality of battery strings are electrically connected in parallel.
9. The method of claim 7 , further comprising generating, with the control system, a warning identifying the determined electrical loading limit.
10. The method of claim 9 , further comprising decreasing, with the control system, an electrical loading of the aircraft electrical distribution bus in response to detection of the fault in a faulted first battery string.
11. An assembly for an aircraft, the assembly including:
an aircraft electrical distribution bus;
an aircraft propulsion system including a rotational assembly, the rotational assembly including a rotatable shaft and an electric motor, the electric motor connected in electrical communication with the aircraft electrical distribution bus and configured to selectively drive the rotatable shaft;
a battery including a plurality of battery strings and a main battery contactor,
each battery string including a string contactor, the string contactor positionable in a closed string position or an open string position, and
the main battery contactor electrically connected in series with the string contactor of each battery string, the main battery contactor positionable in a closed battery position or an open battery position; and
a control system in signal communication with the string contactor of each battery string, the control system configured to control a position of the string contactor of each battery string in the closed string position or the open string position, the control system including a processor in communication with a non-transitory memory storing instructions, which instructions when executed by the processor, cause the processor to:
detect a fault in a faulted first battery string of the plurality of battery strings;
position the string contactor of the faulted first battery string in the open string position, in response to the detected fault in the faulted first battery string, and maintain the string contactor of at least one unfaulted second battery string of the plurality of battery strings in the closed string position; and
determine an electrical loading limit for the at least one unfaulted second battery string with the string contactor of the faulted first battery string in the open string position.
12. The assembly of claim 11 , wherein the instructions, when executed by the processor, further cause the processor to control the electric motor to decrease an electrical loading of the electric motor in response to detection of the fault in the faulted first battery string.
13. The assembly of claim 12 , wherein:
the propulsion system includes a gas turbine engine, the gas turbine engine including the rotational assembly; and
the instructions, when executed by the processor, further cause the processor to control the gas turbine engine to increase an engine power of the gas turbine engine in response to detection of the fault in a faulted first battery string.
14. The assembly of claim 13 , wherein the rotational assembly further includes a propeller, the electric motor configured to selectively drive the propeller via the rotatable shaft.
15. The assembly of claim 13 , wherein the gas turbine engine includes a generator connected in electrical communication with the aircraft electrical distribution bus.
16. The assembly of claim 15 , wherein the gas turbine engine includes a second rotational assembly including a second rotatable shaft, a bladed turbine rotor, and a bladed compressor rotor, the second rotatable shaft interconnecting the bladed turbine rotor and the bladed compressor rotor, the bladed turbine rotor configured to drive the generator via the second rotatable shaft.
17. The assembly of claim 13 , wherein each battery string of the plurality of battery strings includes a plurality of battery modules electrically connected in series and the plurality of battery strings are electrically connected in parallel.