WO2024060567A1 - 电动飞机综合电推进系统 - Google Patents

电动飞机综合电推进系统 Download PDF

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Publication number
WO2024060567A1
WO2024060567A1 PCT/CN2023/085041 CN2023085041W WO2024060567A1 WO 2024060567 A1 WO2024060567 A1 WO 2024060567A1 CN 2023085041 W CN2023085041 W CN 2023085041W WO 2024060567 A1 WO2024060567 A1 WO 2024060567A1
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WO
WIPO (PCT)
Prior art keywords
bus
power supply
electric propulsion
fly
propulsion system
Prior art date
Application number
PCT/CN2023/085041
Other languages
English (en)
French (fr)
Inventor
黄劲松
戴泽华
查振羽
钱仲焱
吴昊
徐州
Original Assignee
中国商用飞机有限责任公司
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Application filed by 中国商用飞机有限责任公司 filed Critical 中国商用飞机有限责任公司
Publication of WO2024060567A1 publication Critical patent/WO2024060567A1/zh

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Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENTS OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D27/00—Arrangement or mounting of power plant in aircraft; Aircraft characterised thereby
    • B64D27/02—Aircraft characterised by the type or position of power plant
    • B64D27/24—Aircraft characterised by the type or position of power plant using steam, electricity, or spring force
    • H—ELECTRICITY
    • H02—GENERATION; CONVERSION OR DISTRIBUTION OF ELECTRIC POWER
    • H02J—CIRCUIT ARRANGEMENTS OR SYSTEMS FOR SUPPLYING OR DISTRIBUTING ELECTRIC POWER; SYSTEMS FOR STORING ELECTRIC ENERGY
    • H02J1/00—Circuit arrangements for dc mains or dc distribution networks
    • H02J1/08—Three-wire systems; Systems having more than three wires
    • H02J1/084—Three-wire systems; Systems having more than three wires for selectively connecting the load or loads to one or several among a plurality of power lines or power sources
    • H02J1/086—Three-wire systems; Systems having more than three wires for selectively connecting the load or loads to one or several among a plurality of power lines or power sources for providing alternative feeding paths between load or loads and source or sources when the main path fails
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present invention relates to electric propulsion technology for aircraft, and more specifically to an integrated electric propulsion system for electric aircraft.
  • the specific energy of the battery also limits the output capability of the distributed electric propulsion system.
  • the aircraft must comprehensively consider the impact of thrust output and flight attitude on flight performance, so as to make full use of limited electrical power and energy resources and improve the aircraft's flight capabilities.
  • the electrified flight control system also known as the fly-by-wire flight control actuation system
  • the electrified flight control system has short-term, high-power charge and discharge characteristics, which results in the weight and loss of its upstream branch power transmission and distribution devices being much higher than those of traditional flight control systems. system.
  • the present invention is made to solve the above-mentioned existing technical problems, and its purpose is to provide an integrated electric propulsion system for electric aircraft that integrates the distributed electric propulsion system and the fly-by-wire flight control actuation system integrated on the wing.
  • Electric aircraft provide reliable, efficient, and stable flight power while reducing the weight and energy loss of the system.
  • the present invention provides an integrated electric propulsion system for an electric aircraft.
  • the integrated electric propulsion system for an electric aircraft is independent of an airborne electric energy system that connects multiple electrified loads. It is characterized in that the integrated electric propulsion system for an electric aircraft is
  • the propulsion system integrates the distributed electric propulsion system and the fly-by-wire flight control actuation system integrated on the wings of the electric aircraft into a unified power supply source.
  • the first power supply path is used to supply the distributed electric propulsion system to the distributed power supply.
  • the propulsion system and the fly-by-wire flight control actuation system provide power, and in the event of a partial failure of the normal power supply or an emergency situation of complete failure, the energy management department will provide power to the distributed electric propulsion system.
  • the power supply path of the fly-by-wire flight control actuation system is switched to a power supply path different from the first path.
  • the electric aircraft integrated electric propulsion system of the present invention has the following beneficial effects:
  • the integrated electric propulsion system of the electric aircraft of the present invention and the airborne electric energy system are independent of each other, have simple control, low implementation difficulty, and are easy to implement and modify;
  • the integrated electric propulsion system of the electric aircraft of the present invention has the characteristics of a distributed energy system, reducing the power supply cables and conversion devices of the equipment, reducing the weight of the system, reducing the loss of electrical energy, and improving the reliability of the system;
  • the electric aircraft integrated electric propulsion system of the present invention integrates the two system functions of thrust output (distributed electric propulsion system) and attitude control (fly-by-wire flight control actuation system), providing a basis for the comprehensive and coordinated control of the two, which is beneficial to improving the performance, stability and energy utilization of the electric aircraft.
  • the electric aircraft integrated electric propulsion system of the present invention adopts a unified power supply to connect the distributed electric propulsion system and the fly-by-wire flight control actuation system.
  • the main power supply that has not failed can be used to power the entire distributed electric propulsion system and the entire fly-by-wire flight control actuation system through a power supply path different from the first power supply path during normal power supply, thereby maximizing the thrust output and attitude control system functions of the electric aircraft.
  • the backup power supply that has not failed can still be used to power the entire fly-by-wire flight control actuation system, thereby ensuring the safety of the electric aircraft through attitude control. Landing.
  • the energy management unit is a plurality of switchboard boxes.
  • Multiple switchboard boxes can be installed symmetrically on the left and right wings, or the number of switchboard boxes installed on one side of the wing is different from the number of switchboard boxes installed on one side of the wing. Number of switchboard boxes.
  • each wing on each side includes a distribution panel box.
  • each distribution panel box includes a plurality of first DC busbars connected in series and a plurality of second DC busbars connected in series.
  • the first DC busbars electrically adjacent to each other are respectively connected to each other through a first branch; among the plurality of second DC busbars connected in series, the second DC busbars electrically adjacent to each other are respectively connected to each other through a second branch; and the plurality of first DC busbars and the plurality of second DC busbars are respectively connected to each other through a third branch provided with a voltage conversion device.
  • one or more first DC busbars of the distribution panel box on one side are correspondingly connected to one or more first DC busbars of the distribution panel box on the other side through electrical lines
  • one or more second DC busbars of the distribution panel box on one side are correspondingly connected to one or more second DC busbars of the distribution panel box on the other side through electrical lines, thereby forming the energy management unit.
  • the wing it is also possible to include more than one switchboard box on one side of the wing and more than two switchboard boxes on the other side of the wing.
  • the three or more distribution board boxes on both sides it is only necessary that at least one of the distribution board boxes has the structure of the distribution board box of the above example, and the first distribution box is connected between two or more three distribution boxes through electrical lines.
  • the DC bus bars are connected to each other, and the second DC bus bars are connected to each other through another electrical line, thereby forming an energy management part.
  • first DC bus bar in each switchboard box is connected to the first battery The system is connected, and the second DC bus bar in each distribution panel box is connected to the second battery system, and the first battery system and the second battery system form a unified power supply source.
  • each first DC bus bar of the plurality of first DC bus bars allows selection of one by the first battery system and the first DC bus bar electrically adjacent to the first DC bus bar. power supply path.
  • each of the plurality of second DC bus bars allows for a second DC bus system to be connected by a second battery system, a first DC bus bar electrically adjacent to the second DC bus bar, and a first DC bus bar connected to the second DC bus bar. A supply path is selected by electrically adjacent second DC bus bars.
  • each switchboard box of the electric aircraft integrated electric propulsion system has multiple power channels with the first battery system as the main power source, and when an emergency occurs, multiple second battery systems serve as backup power sources.
  • a fly-by-wire actuating unit on a second DC bus provides emergency power.
  • each of the plurality of first DC bus bars is respectively powered by a directly connected first battery system, and each of the plurality of second DC bus bars is also respectively powered by a directly connected voltage converter.
  • the equipment receives power from the electrically adjacent connected first DC bus. Therefore, when supplying normal power to each first DC bus, the first power supply path to supply power to the distributed electric propulsion system and the fly-by-wire flight control actuation system is: through each set of first battery systems to the corresponding power supply in the distribution board box.
  • the first DC bus provides power to ensure the operation of the electric propulsion units that constitute the distributed electric propulsion system connected to each of the first DC buss. Then the current will flow into the corresponding third branch through the corresponding third branch.
  • Each of the second DC bus bars electrically connected to the first DC bus bar is connected to each of the second DC bus bars to ensure that the fly-by-wire flight control actuation unit that constitutes the fly-by-wire actuation system is connected to each of the second DC bus bars. of operation.
  • the first DC bus bar corresponding to the first battery system in the power supply channel is switched to be powered by an adjacent power supply channel. Therefore, when a partial failure occurs in the normal power supply to each first DC bus, that is, when a partial failure occurs in the normal power supply branches of multiple sets of first battery systems, the power supply to the distributed electric propulsion system and fly-by-wire flight control
  • the second power supply path of the actuating system is: through the adjacent first battery system that has not failed to the corresponding first DC sink in the distribution board box.
  • the flow bar supplies power to ensure the operation of the electric propulsion unit that constitutes the distributed electric propulsion system connected to the first DC bus, and then the current flows into the electrical phase of the first DC bus through the corresponding third branch.
  • the second DC bus connected adjacent to the second DC bus ensures the operation of the fly-by-wire actuating unit that constitutes the fly-by-wire actuating system. At the same time, the current will flow from the first battery that has not failed.
  • the first DC bus corresponding to the system supplies power through the first branch to the first DC bus corresponding to the failed first battery system, that is, the first DC bus that is the redundant power supply object, so as to ensure that it is connected to the first DC bus that is the target of redundant power supply.
  • the bus bar is electrically connected to a second DC bus bar that is electrically adjacent to it, or flows from the second DC bus bar that is electrically adjacent to it and flows into the first DC bus bar that is the redundant power supply object through the corresponding second branch.
  • a second DC bus bar electrically connected adjacently.
  • first DC bus bar is not completely electrically disconnected from the first battery system and the first DC bus electrically connected to the first DC bus bar, it can be ensured that the first DC bus bar is connected to the first DC bus bar.
  • the operation of the electric propulsion unit connected to the first DC bus bar can maximize the two system functions of thrust output and attitude control of the electric aircraft.
  • each of the plurality of second DC bus bars has a redundant power supply capability that can be powered by a plurality of second battery systems.
  • the fourth power supply path to the fly-by-wire flight control actuation system is: through the adjacent non-failed second battery system to the switchboard box
  • the corresponding second DC bus supplies power, and at the same time, the current will flow from the second DC bus corresponding to the non-failed second battery system to the second DC bus corresponding to the failed second battery system via the corresponding second branch.
  • the second DC bus bar that is the redundant power supply object supplies power to the maximum Ensure the operation of the fly-by-wire flight control actuation system to the maximum extent possible.
  • the integrated electric propulsion system of the electric aircraft is installed symmetrically on the left and right wings of the electric aircraft, and each wing on each side includes a distribution panel box, Four sets of electric propulsion units constituting the distributed electric propulsion system, four sets of fly-by-wire flight control actuation units constituting the fly-by-wire flight control actuation system, and two sets of first battery systems and two sets of second batteries as unified power supply sources system.
  • a distribution panel box Four sets of electric propulsion units constituting the distributed electric propulsion system, four sets of fly-by-wire flight control actuation units constituting the fly-by-wire flight control actuation system, and two sets of first battery systems and two sets of second batteries as unified power supply sources system.
  • the first battery system is, for example, a high specific energy lithium battery system.
  • the second battery system is, for example, an emergency lithium battery system.
  • the second battery system (emergency lithium battery system) may be a battery system with a charge and discharge function, or a battery system with only a discharge function. In the former case, it is preferable to have a DC/DC converter capable of charging and discharging.
  • Figure 1 is a module schematic diagram of the integrated electric propulsion system of the electric aircraft of the present invention.
  • Figure 2 is a schematic diagram of the physical structure of the electric aircraft integrated electric propulsion system of the present invention arranged on the electric aircraft.
  • Figure 3 is a system architecture diagram of the integrated electric propulsion system of the electric aircraft of the present invention arranged on the electric aircraft.
  • Figure 4 is a system architecture diagram of the wing located on one side of the integrated electric propulsion system of the electric aircraft of the present invention, which shows the specific structure inside the switchboard box.
  • CA CA, CB electrical wiring.
  • the electric aircraft integrated electric propulsion system 100 of the present invention is independent of the electrified airborne system that connects multiple electrified loads, that is, the airborne electric energy system. As shown in Figure 1, it is a distribution that will be integrated on the wing 11 of the electric aircraft 10.
  • the electric propulsion system 130 and the fly-by-wire flight control actuation system 140 are integrated and connected to a unified power supply 110, and provide a unified and mutually coordinated electric energy supply for the flight of the electric aircraft through the energy management unit 120, while effectively reducing the Weight and energy consumption of electric aircraft.
  • the energy management unit 120 manages, distributes and transmits the electric energy of the unified power supply 110 through a plurality of switchboard boxes 121 described later, thereby supplying power to the distributed electric propulsion system 130 and the fly-by-wire flight control actuation system 140 .
  • the distributed electric propulsion system 130 uses multiple sets of electric propulsion units 131 described below as thrust generating devices to provide thrust for aircraft flight.
  • the fly-by-wire actuation system 140 uses multiple sets of fly-by-wire actuation units 141 described below to change the inclination angles of ailerons, flaps, etc., to provide different flight postures for the aircraft.
  • Figure 2 shows the physical structure of the electric aircraft integrated electric propulsion system 100 of the present invention arranged on the electric aircraft 10.
  • Figure 3 is a system architecture diagram of the electric aircraft integrated electric propulsion system 100 of the present invention arranged on the electric aircraft 10. .
  • the integrated electric propulsion system 100 of the electric aircraft of the present invention is installed symmetrically on the left and right wings 11 of the electric aircraft 10 , and on each side
  • Each wing 11 includes at least one (one shown in Figure 2) switchboard box 121, multiple sets (four sets shown in Figure 2) of electric propulsion units 131 constituting a distributed electric propulsion system 130, and a fly-by-wire flight system.
  • Multiple sets of control actuation systems 140 (shown in Figure 2 as Four sets) fly-by-wire actuating units 141.
  • FIG. 3 multiple sets of electric propulsion units 131 and multiple sets of fly-by-wire actuating units 141 located on the wings 11 on each side are respectively located on that side.
  • the switchboard box 121 is connected.
  • Each set of electric propulsion unit 131 includes a high specific power three-phase inverter 131a, a high specific power motor 131b and a propulsion device 131c.
  • the high specific power three-phase inverter 131a controls the thrust output by the electric propulsion unit 131.
  • the high specific power motor 131b converts three-phase electrical energy into mechanical speed and torque.
  • the propulsion device 131c uses a propeller or a ducted fan to convert the rotation speed and torque output by the motor into thrust.
  • fly-by-wire actuating units 141 are symmetrically arranged on the left and right sides of the rear edge of the wing 11 and are powered by second (for example, 28V low voltage) DC bus bars 121L in the respective distribution panel boxes 121 .
  • Each fly-by-wire actuator unit 141 includes a three-phase inverter 141a, a servo motor 141b and an actuator 141c.
  • the servo motor 141b outputs a certain rotation speed according to the control command of the three-phase inverter 141a.
  • the actuator 141c is controlled to deflect the control surfaces of the ailerons, flaps, etc., thereby adjusting the flight attitude.
  • the fly-by-wire flight control actuating unit 141 is powered by a 28V low-voltage DC bus, for example.
  • the voltage level of the power supply bus of the fly-by-wire flight control actuator unit 141 will increase.
  • a 270V high-voltage DC bus is used for power supply. This depends on the specific size of the electric aircraft 10 and the voltage level of the fly-by-wire flight control actuator unit 141. power.
  • each wing 11 on each side also includes multiple sets (two sets are shown in FIG. 2 ) of first battery systems 111 and multiple sets (two sets are shown in FIG. 2 ) of second battery systems 112.
  • the multiple sets of first battery systems 111 and the multiple sets of second battery systems 112 on each wing 11 are respectively connected to the switchboard box 121.
  • the first battery systems 111 and the second battery systems 112 on the left and right wings 11 are installed in a left-right symmetrical manner to form a unified power supply 110 shown in FIG. 1 , which is used to supply power to the switchboard boxes 121 on the left and right sides.
  • the first battery system 111 is, for example, a high specific energy lithium battery system, which includes a lithium battery pack 111 a and a power conversion device 111 b .
  • the lithium battery pack 111a is composed of several lithium battery cells and two independent battery management systems, thermal management systems, shells, relays, sensors, etc.
  • Each lithium battery cell is composed of one or more lithium battery cells.
  • the two battery management systems monitor the status of the lithium battery cells in different ways and have communication and control functions.
  • the thermal management system can dynamically adjust the temperature of the lithium battery pack 111a to ensure that the operating temperature of the lithium battery pack 111a is within an appropriate range.
  • the lithium battery shell has fireproof and explosion-proof functions.
  • the relay can open and cut off the high-voltage electrical circuit of the lithium battery cell to achieve external on and off.
  • the power conversion device 111b is a DC/DC converter, which can control charge and discharge according to the status of the lithium battery pack and flight power requirements.
  • the second battery system 112 is, for example, an emergency lithium battery system, which has a lithium battery pack 112a and a DC/DC converter 112b that can charge and discharge, and is used for key equipment of the integrated electric propulsion system in emergency situations. powered by.
  • the lithium battery pack 112a can be charged by switching the DC/DC converter 112b to the charging mode, and in the second battery system 112 performs trickle charging when it is nearly fully charged.
  • the DC/DC converter 112b can be switched to the discharge mode to output the power of the lithium battery pack 112a.
  • the electric aircraft integrated electric propulsion system 100 uses at least one first DC bus bar 121H and at least one second DC bus bar respectively in the distribution panel box 121 located on the left and right wings 11 121L are connected through electrical lines CA and CB respectively, and the switchboard boxes 121 located on the left and right wings 11 constitute the pair of power supply sources 110 (multiple sets of first battery systems 111 and multiple sets of second batteries) shown in Figure 1
  • the energy management unit 120 manages, distributes and transmits the electric energy of the system 112), so that the unified power supply 110 can be used to control the corresponding sets of electric propulsion units 131 and each set of electric wires through the control of the switchboard box 121 on the corresponding side.
  • the power supply of the flight control actuation unit 141 thus provides a multi-channel redundant
  • the additional backup power supply channel improves the fault tolerance performance of the power supply 110.
  • the electric propulsion units 131 located on the left and right wings 11 constitute the distributed electric propulsion system 130 shown in FIG. 1 .
  • the fly-by-wire actuation units 141 located on the left and right wings 11 constitute the fly-by-wire actuation system 140 shown in FIG. 1 .
  • Figure 4 is a system architecture diagram of the wing 11 located on one side of the electric aircraft integrated electric propulsion system 10 of the present invention, which shows the specific internal structure of the switchboard box 121.
  • each distribution panel box 121 includes a plurality (two in FIG. 4 ) of the first DC bus bars 121H connected in series and a plurality (two in FIG. 4 ) of the first DC bus bars 121H connected in series.
  • the second DC bus bars 121L and a plurality of series-connected first DC bus bars 121H two electrically adjacent first DC bus bars 121H are connected to each other through the first branch C1 respectively, and a plurality of series-connected first DC bus bars 121H Among the connected second DC bus bars 121L, two electrically adjacent second DC bus bars 121L are connected to each other through the second branch C2.
  • first DC bus bars 121H and a plurality of second DC bus bars 121H are connected to each other through the second branch C2.
  • the DC bus bars 121L are connected to each other one by one through the third branch C3 provided with the voltage conversion device 121a.
  • the voltage conversion device 121a is, for example, a DC/DC converter, which converts the DC power at the voltage provided by the first DC bus bar 121H into the DC power at the voltage required by the second DC bus bar 121L.
  • first battery systems 111 are respectively connected to multiple first DC bus bars 121H in each distribution panel box 121 , and each first DC bus bar 121H in each distribution panel box 121 is connected to multiple sets of first DC bus bars 121H in each distribution panel box 121 .
  • electric propulsion units 131 are connected.
  • multiple sets of second battery systems 112 are respectively connected to a plurality of second DC bus bars 121L in each distribution panel box 121 , and each second DC bus bar 121L in each distribution panel box 121 is connected to a plurality of sets (Fig. Two sets are shown in 4) fly-by-wire actuating units 141 are connected.
  • each switchboard box 121 also includes switching devices such as bus bar contactors, control relays, and thermal circuit breakers.
  • each branch circuit (the first branch circuit C1, the second branch circuit C2 and the third branch circuit C3) is equipped with line switching protection equipment such as contactors or circuit breakers.
  • a redundant and fault-tolerant power grid structure is used inside the distribution panel box 121 to ensure the safe and reliable operation of the power grid of the electric aircraft integrated electric propulsion system 100 .
  • each set of the first battery system 111 supplies power to the corresponding first DC bus 121H in the switchboard box 121 to ensure the operation of the electric propulsion unit 131 connected to each first DC bus 121H, and then the current will flow through the corresponding third branch C3 into the second DC bus 121L electrically adjacent to each first DC bus to ensure the operation of the fly-by-wire flight control actuation unit 141 connected to each second DC bus 121L.
  • the power supply path at this time i.e., the first power supply path, is That is, under normal power supply conditions, each first DC busbar 121H is directly powered by each first battery system 111 , and each second DC busbar 121L is powered by each first DC busbar 121H electrically connected thereto via the voltage conversion device 121 a .
  • the plurality of first DC bus bars 121H have redundant power supply capabilities.
  • the first DC bus bar 121H corresponding to a battery system 111 can operate normally and is powered by the adjacent first battery system 111 that has not failed. That is, the adjacent first battery system 111 that has not failed first supplies power to the distribution board box 121
  • the first DC bus 121H corresponding to the non-failed first battery system 111 supplies power to ensure the operation of the electric propulsion unit 131 corresponding to the first DC bus 121H, and then the current will pass through the corresponding third branch.
  • Path C3 flows into the second DC bus 121L electrically connected to the first DC bus 121H to ensure the operation of the fly-by-wire flight control actuating unit 141 connected to the second DC bus 121L; at the same time, , the current will flow from the first DC bus bar 121H corresponding to the non-failed first battery system 111 to the first DC bus bar 121H corresponding to the failed first battery system 111 via the corresponding first branch C1 (also (called “the first DC bus as the redundantly powered object”) to ensure power supply to the first DC bus 121H (i.e., the first DC bus 121H as the redundantly powered object).
  • the electric propulsion unit 131 operates next if the first DC bus 121H as the redundant power supply object and the second DC bus 121L electrically connected adjacently (also known as the "second DC bus as the power supply object" ) is the path (the circuit breaker is closed ), the current will flow from the first DC bus 121H as the redundant power supply object through the corresponding third branch C3 into the second DC bus 121L as the power supply object.
  • the third branch C3 between the first DC bus 121H and the second DC bus 121L used as the power supply object is open circuit (the circuit breaker is opened), then the current cannot flow from the corresponding third branch C3 to the power supply object
  • the second DC bus 121L electrically connected to the second DC bus 121L serving as the power supply target is connected via the corresponding second branch C2 (this second DC bus
  • the corresponding third branch C3 between 121L and the corresponding first DC bus 121H supplies the path (circuit breaker closed)).
  • the second power supply path is switched to a second power supply path different from the first power supply path, and the first DC bus bar 121H that cannot obtain power from the failed first battery system 111 can pass through other power supply paths.
  • the first DC bus bar 121H can obtain power indirectly from the first DC bus bar 121H that normally obtains power from the failed first battery system 111, and the second DC bus bar 121L can obtain power according to the first DC bus bar 121H that is electrically adjacently connected.
  • the third branch C3 between them is connected and powered through the first DC bus 121H or other second DC bus 121L connected electrically adjacent thereto.
  • the two-battery system 112 supplies power to the corresponding second DC bus 121L to ensure the operation of the fly-by-wire flight control actuating unit 141 connected to each second DC bus 121L.
  • the power supply branch that is, the emergency power supply branch is
  • multiple second DC bus bars 121L also have redundant power supply capabilities.
  • the DC bus bar 121H can operate normally and supply power to it through the adjacent second battery system 112 that has not failed. That is, the adjacent second battery system 112 that has not failed first supplies power to the corresponding second DC bus in the distribution panel box 121. Bar 121L supplies power; at the same time, the current will flow from the second DC bus 121L corresponding to the non-failed second battery system 112 to the second DC bus corresponding to the failed second battery system 112 via the corresponding second branch C2.
  • the second DC bus bar 121L (also known as "the second DC bus as the redundant power supply object") supplies power to ensure the operation of the fly-by-wire flight control actuation system 140 to the maximum extent. That is, when all normal power supply branches fail, switching to a third power supply path or a fourth power supply path that is different from the first power supply path and the second power supply path, the second DC bus bar 121L can be connected via the corresponding second battery. The system 112 or other second DC bus 121L electrically connected thereto is powered.
  • each switchboard box 121 of the electric aircraft integrated electric propulsion system 100 of the present invention has a plurality of power channels using the first battery system 111 as the power supply.
  • the plurality of first DC bus bars Each of the plurality of second DC bus bars 121H is respectively powered by a respective directly connected first battery system 111, and each of the plurality of second DC bus bars 121L is also respectively powered from an electrically adjacent connected first DC bus bar via a directly connected voltage conversion device 121a. 121H obtains power.
  • the first DC bus bar 121H corresponding to the first battery system 111 in the power supply channel will switch to be powered by an adjacent power supply channel.
  • the plurality of second battery systems 112 provide emergency power for the fly-by-wire actuating units 141 on the plurality of second DC bus bars 121L, and each of the plurality of second DC bus bars 121L is specifically The redundant power supply capability can be powered by multiple second battery systems 112 .
  • each first DC bus 121H (also called a “target first DC bus”) among the plurality of first DC buses 121H is allowed to be connected to the target first DC bus.
  • the directly connected first battery system 111 and the first DC bus bar electrically adjacent to the first DC bus bar 121H (ie, the target first DC bus bar) select a power supply path.
  • Each second DC bus 121L of the plurality of second DC buses 121L (also referred to as a "target second DC bus”) allows the second battery system 112 to be directly connected to the target second DC bus.
  • the first DC bus bar electrically connected to the second DC bus bar 121 ie, the target second DC bus bar
  • the first DC bus bar electrically connected to the second DC bus bar 121 ie, the target second DC bus bar
  • Adjacently connected second DC busbars select a supply path.
  • first DC bus or the target second DC bus If the electrically adjacent first DC bus bar and/or the second DC bus bar cannot be directly connected (power supplied) from the first battery system 111 or the second battery system 112, then the next electrically adjacent electrically adjacent DC bus bar is selected again. Connect the first DC bus and/or the second DC bus.
  • the switchboard box 121, the first battery system 111, the second battery system 112, the electric propulsion unit 131, the fly-by-wire flight control actuating unit 141, and the first DC bus 121H , the second DC bus bar 121L, etc. are not limited to the number shown in FIG. 2 .
  • the electric aircraft integrated electric propulsion system 100 of the present invention it is not limited to only one switchboard box 121 being provided on each of the left and right wings 11 , but more than two switchboard boxes 121 can also be provided respectively. .
  • the number of switchboard boxes 121 disposed on the left and right wings 11 is preferably the same, but may also be different.
  • one switchboard box 121 may be disposed on one wing 11 , and two switchboard boxes 121 are provided on the wing 11 on the other side.
  • each first battery system 111 corresponds to more than two sets of electric propulsion units 131
  • Each set of second battery systems 112 can correspond to more than two sets of fly-by-wire flight control actuating units 141 .
  • each switchboard box 121 has the aforementioned structure.
  • first DC bus bars 121H and two second DC bus bars 121L which are configured as: one (for example, the upper left one in Figure 4)
  • the first DC bus bar 121H is electrically adjacent (e.g., Figure 4
  • the other first DC bus bar 121H on the upper right side in Figure 4 is connected through the first branch C1;
  • One second DC bus bar 121L at the lower left in Figure 4 is connected through the third branch C3;
  • the other second DC bus 121L at the bottom is connected through the second branch C2;
  • the other (for example, the upper right in Figure 4) first DC bus 121H is electrically adjacent (for example, the right in Figure 4
  • the lower) second DC bus bar 121L is connected through the third branch C3, but the structure of the distribution panel box 121 of the electric aircraft integrated electric propulsion system 100 of the present invention can also have more than three first DC bus bars 121H and More than three second DC bus bars 121L, taking the case of three as an example, they can
  • the other first DC bus bar 121H is electrically adjacent to another first DC bus bar 121H (for example, the one on the upper right).
  • the first DC bus bar 121H is also connected through the first branch C1.
  • the circuit C1 is connected; one (for example, the lower left) second DC bus 121L is connected to another electrically adjacent (for example, the middle lower) second DC bus 121L through the second branch C2, and the other second DC bus 121L is connected through the second branch C2.
  • the bus bar 121L is also connected to a second DC bus bar 121L electrically adjacent (for example, the one on the lower right) through the second branch C2; a first DC bus 121H (for example, the one on the upper left) is electrically adjacent to the One (for example, the lower left) second DC bus 121L is connected through the third branch C3, and the other (for example, the upper middle) first DC bus 121H is electrically adjacent to another (for example, the lower middle) first DC bus 121H.
  • Two DC bus bars 121L are connected through the third branch C3, and another first DC bus bar 121H (for example, the one on the upper right) is electrically adjacent to another second DC bus bar 121L (for example, the one on the lower right) through the third branch C3.
  • Branch C3 is connected.
  • a component C other than component A and component B.
  • a voltage conversion device 121a and a circuit breaker
  • the electrically adjacent components A and B do not necessarily refer to two components that are physically adjacent to each other or very close to each other, but may also be two components that are physically far apart, as long as the above conditions are met.
  • the distribution board box 121 on the wing 11 on the left side (for example, the upper side in Figure 3) is connected to the switchboard box 121 on the right side.
  • the distribution board box 121 on the wing 11 on one side (for example, the lower side in FIG. 3 ) is configured such that only one of the first DC bus bars 121H is connected through the electrical line CA and only one of the second DC bus bars 121L is connected thereto.
  • the electrical lines CB are electrically connected to each other, but the present invention is not limited thereto. It can also be configured such that each first DC bus bar 121H is connected correspondingly through the electrical line CA and each second DC bus bar 121H is connected to each other. 121L is correspondingly connected via another electrical line CB.
  • the first battery system 111 is a high specific energy lithium battery system and the second battery system 112 is an emergency lithium battery system.
  • the first battery system 111 And the second battery system 112 may be any battery system suitable for use in electric aircraft.
  • the electric capacity of the first battery system 111 is greater than the electric capacity of the second battery system 112 .
  • the second battery system 112 is exemplified as a battery system that can be charged and discharged, the second battery system 112 does not necessarily have a charging function, and may also be a battery system that only discharges.

Abstract

一种电动飞机综合电推进系统，能整合集成在机翼上的分布式电推进系统和电传飞控作动系统，为电动飞机提供可靠、高效、稳定的飞行动力，同时降低系统的重量和能量损耗。所述电动飞机综合电推进系统独立于挂接多个电气化负载的机载电能系统，所述电动飞机综合电推进系统将集成在电动飞机的机翼上的分布式电推进系统和电传飞控作动系统综合成与统一的供电源连接，在正常供电时，以第一供电路径向所述分布式电推进系统和所述电传飞控作动系统供电，并且在所述正常供电发生部分失效的故障情况时或是发生全部失效的应急情况时，通过能量管理部，将向所述分布式电推进系统和所述电传飞控作动系统供电的路径切换成与所述第一供电路径不同的供电路径。

Description

电动飞机综合电推进系统 技术领域
本发明涉及飞机的电推进技术，更具体地涉及一种电动飞机综合电推进系统。
背景技术
随着电气化技术的发展，电动飞机采用电池和推进电机替代传统飞机的燃油和发动机的驱动方式，在降低噪声和排放的同时提高了能源利用率。目前受到电机技术和配电技术的限制，通行做法是将多个电推进单元布置在飞机机翼边沿，集成为分布式电推进系统，从而在满足已有成熟电力电子技术的条件下，为电动飞机提供有效推力。
但是，机翼各处的推力产生的力矩不同，若仅仅采用分布式电推进系统，则可能会对飞机飞行的稳定性造成巨大的负面影响。
同时，电池比能量也限制了分布式电推进系统的输出能力，飞机必须综合考虑推力输出和飞行姿态对飞行性能的影响，从而充分利用有限的电功率和能量资源，提高飞机的飞行能力。
此外，电气化飞控系统(又称电传飞控作动系统)具有短时、大功率充放电的特性，这导致其上游支路的输电、配电装置重量和损耗远远高于传统飞控系统。
发明内容
因此，本发明为解决上述现有各技术问题而作，其目的在于提供一种电动飞机综合电推进系统，整合集成在机翼上的分布式电推进系统和电传飞控作动系统，为电动飞机提供可靠、高效、稳定的飞行动力，同时降低系统的重量和能量损耗。
为了实现上述目的，本发明提供了一种电动飞机综合电推进系统，所述电动飞机综合电推进系统独立于挂接多个电气化负载的机载电能系统，其特征是，所述电动飞机综合电推进系统将集成在电动飞机的机翼上的分布式电推进系统和电传飞控作动系统综合成与统一的供电源连接，在正常供电时，以第一供电路径向所述分布式电推进系统和所述电传飞控作动系统供电，并且在所述正常供电发生部分失效的故障情况或是发生全部失效的应急情况时，通过能量管理部，将向所述分布式电推进系统和所述电传飞控作动系统供电的路径切换成与所述第一路径不同不同的供电路径。
根据如上所述构成，相比于现有技术的架构，本发明的电动飞机综合电推进系统具有以下有益效果：
(1)本发明的电动飞机综合电推进系统与机载电能系统互相独立，控制简单、实现难度低，易于实现和改装；
(2)本发明的电动飞机综合电推进系统具有分布式能源系统的特点，减少了设备供电的线缆以及转换装置，减轻了系统重量，降低了电能的损耗，也提高了系统的可靠性；
(3)本发明的电动飞机综合电推进系统集成了推力输出(分布式电推进系统)和姿态控制(电传飞控作动系统)两个系统功能，为两者的综合、协同控制提供基础，有利于提升电动飞机的性能、稳定性和能量利用率。
(4)本发明的电动飞机综合电推进系统通过采用统一的供电源与分布式电推进系统和电传飞控作动系统连接，在一部分主电源发生故障(失效)，能够利用未失效的主电源，以与正常供电时的第一供电路径不同的供电路径，对整套分布式电推进系统和整套电传飞控作动系统进行供电，最大限度保障电动飞机的推力输出和姿态控制两个系统功能。即使是主电源全部发生故障(失效)，甚至是一部分备用电源也发生故障(失效)，仍能够利用未失效的备用电源对整套电传飞控作动系统进行供电，从而能通过姿态控制确保电动飞机的安全 着陆。
优选的是，所述能量管理部是多个配电盘箱。多个配电盘箱既可以呈左、右对称地安装于左、右两侧的机翼，也可以是安装在一侧的机翼上的配电盘箱的数量不同于安装在一侧的机翼上的配电盘箱的数量。
在一个实例中，例如，在每一侧的所述机翼上各自包括一个配电盘箱。并且，较佳地，每个所述配电盘箱包括多个串联连接的第一直流汇流条和多个串联连接的第二直流汇流条。在这种结构的配电盘箱中，多个串联连接的第一直流汇流条中，两两电气相邻的所述第一直流汇流条之间分别通过第一支路相互连接；多个串联连接的第二直流汇流条中，两两电气相邻的所述第二直流汇流条之间分别通过第二支路相互连接；多个所述第一直流汇流条与多个所述第二直流汇流条逐个对应地通过设置有电压转换设备的第三支路相互连接。多个配电箱盘中，一侧的配电盘箱的一个或多个第一直流汇流条对应地与另一侧的配电盘箱的一个或多个第一直流汇流条通过电气线路相互连接，并且一侧的配电盘箱的一个或多个第二直流汇流条对应地与另一侧的配电盘箱的一个或多个第二直流汇流条通过电气线路相互连接，由此构成所述能量管理部。
由于多个配电盘箱互相之间同样可以设计有多路冗余备份的供电通道，因此，能够提高供电源的电力供给的容错性能。
除了该一个实例之外，也可以是在一侧的机翼上包括一个以上配电盘箱，在另一侧的机翼上包括两个以上配电盘箱。在两侧的三个以上配电盘箱中，只要至少一个所述配电盘箱具有上述一个实例的配电盘箱的结构即可，并将三个以上配电箱两两之间通过电气线路将所述第一直流汇流条相互连接，并通过另一电气线路将所述第二直流汇流条相互连接，由此构成能量管理部。
进一步优选地，各个配电盘箱中的第一直流汇流条与第一电池 系统连接，并且各个配电盘箱中的第二直流汇流条与第二电池系统连接，并且使第一电池系统和第二电池系统构成统一的供电源。
换句话说，多个第一直流汇流条中的每一个第一直流汇流条允许由第一电池系统和与该第一直流汇流条电气相邻连接的第一直流汇流条选择一条供电路径。并且，多个第二直流汇流条中的每一个第二直流汇流条允许由第二电池系统、与该第二直流汇流条电气相邻连接的第一直流汇流条以及与该第二直流汇流条电气相邻连接的第二直流汇流条选择一条供电路径。
根据如上所述构成，电动飞机综合电推进系统的各个配电盘箱内部具有多个以第一电池系统为主电源的通电通道，并且当发生应急情况时，多个第二电池系统作为备用电源为多个第二直流汇流条上的电传飞控作动单元提供应急电源。
在正常供电的情况下，多个第一直流汇流条中的每一个分别由各自直接连接的第一电池系统供电，多个第二直流汇流条中的每一个也分别经由直接连接的电压转换设备从电气相邻连接的第一直流汇流条获得电力。因此，在向各个第一直流汇流条正常供电时，向分布式电推进系统和电传飞控作动系统供电的第一供电路径是：通过每套第一电池系统向配电盘箱中的对应的第一直流汇流条供电，以确保与各所述第一直流汇流条连接的构成分布式电推进系统的电推进单元的运行，接着电流会经由对应的第三支路流入与各所述第一直流汇流条电气相邻连接的各所述第二直流汇流条，以确保与各所述第二直流汇流条连接的构成电传飞控作动系统的电传飞控作动单元的运行。
另外，当某个供电通道中出现故障的故障情况时，与该通电通道中的第一电池系统对应的第一直流汇流条切换至由邻近的通电通道供电。因此，在向各个第一直流汇流条正常供电发生部分失效时，即在多套第一电池系统的正常供电支路存在部分失效的故障情况下，向分布式电推进系统和电传飞控作动系统供电的第二供电路径是：通过邻近的未失效的第一电池系统向配电盘箱中的与其对应的第一直流汇 流条供电，以确保与该第一直流汇流条连接的构成分布式电推进系统的电推进单元的运行，接着电流会经由对应的第三支路流入与该第一直流汇流条电气相邻连接的第二直流汇流条，以确保与该第二直流汇流条连接的构成电传飞控作动系统的电传飞控作动单元的运行，同时电流会从与未失效的第一电池系统对应的第一直流汇流条经由第一支路向与失效的第一电池系统对应的第一直流汇流条即作为被冗余供电对象的第一直流汇流条供电，以确保与作为被冗余供电对象的第一直流汇流条连接的电推进单元的运行，接着电流会从作为被冗余供电对象的第一直流汇流条经由对应的第三支路流入与该第一直流汇流条电气相邻连接的第二直流汇流条，或是从电气相邻连接的第二直流汇流条经由对应的第二支路流入与作为所述被冗余供电对象的第一直流汇流条电气相邻连接的第二直流汇流条。
也就是说，只要某个第一直流汇流条没有与第一电池系统以及与和该第一直流汇流条电气相邻连接的第一直流汇流条完全电气断开，则能确保与该第一直流汇流条连接的电推进单元的运行，由此，能最大限度保障电动飞机的推力输出和姿态控制两个系统功能。
在向各个第一直流汇流条正常供电发生全部失效时，即在多套第一电池系统的正常供电支路全部失效的应急情况下，向电传飞控作动系统供电的第三供电路径是：通过第二电池系统向对应的第二直流汇流条供电，以确保与各第二直流汇流条连接的构成电传飞控作动系统的电传飞控作动单元的运行。
优选的是，多个第二直流汇流条中的每一个具体冗余供电的能力，能由多个第二电池系统供电。此时，当多套第二电池系统的应急供电支路存在部分失效时，向电传飞控作动系统供电的第四供电路径是：通过邻近的未失效的第二电池系统向配电盘箱中的与其对应的第二直流汇流条供电，同时电流会从与未失效的第二电池系统对应的第二直流汇流条经由对应的第二支路向与失效的第二电池系统对应的第二直流汇流条即作为被冗余供电对象的第二直流汇流条供电，以最大 限度确保电传飞控作动系统的运行。
也就是说，只要某个第二直流汇流条没有与第二电池系统、与和该第二直流汇流条电气相邻连接的第一直流汇流条以及与和该第二直流汇流条电气相邻连接的第二直流汇流条完全电气断开，则能确保与该第二直流汇流条连接的电传飞控作动单元的运行。由此，能最大限度保障电动飞机的姿态控制两个系统功能，减少应急情况下电动飞机失控的风险。
在本发明的一个实施例中，电动飞机综合电推进系统呈左、右对称地安装在电动飞机的左、右两侧的机翼上，在每一侧的机翼上各自包括一个配电盘箱、构成分布式电推进系统的四套电推进单元、构成电传飞控作动系统的四套电传飞控作动单元以及作为统一的供电源的两套第一电池系统和两套第二电池系统。但是，应当知道这只是针对某一款或是某一类型/尺寸的电动飞机的特定架构，从更宽泛的意义上说，本发明的具体数量不应当局限于这种特定架构。
所述第一电池系统例如是高比能量锂电池系统。另外，所述第二电池系统例如是应急锂电池系统。
同时，所述第二电池系统(应急锂电池系统)可以是具备充放电功能的电池系统，也可以是仅具备放电功能的电池系统。在前者的情况下，优选的是，具有能充放电的DC/DC变换器。
附图说明
图1是本发明的电动飞机综合电推进系统的模块示意图。
图2是布置在电动飞机上的本发明的电动飞机综合电推进系统的物理结构示意图。
图3是布置在电动飞机上的本发明的电动飞机综合电推进系统的系统架构图。
图4是本发明的电动飞机综合电推进系统的、位于一侧的机翼的系统架构图，其中，示出了配电盘箱内部的具体构成。
(符号说明)
10   电动飞机；
11   机翼；
100  电动飞机综合电推进系统；
110  供电源；
111  第一电池系统；
111a 锂电池包；
111b 电源变换设备；
112  第二电池系统；
112a 锂电池包；
112b DC/DC变换器；
120  能量管理部；
121  配电盘箱；
121a 电压转换设备；
121H 第一直流汇流条；
121L 第二直流汇流条；
130  分布式电推进系统；
131  电推进单元；
131a 高比功率三相逆变器；
131b 高比功率电机；
131c 推进装置；
140  电传飞控作动系统；
141  电传飞控作动单元；
141a 三相逆变器；
141b 伺服电机；
141c 作动筒；
C1   第一支路；
C2   第二支路；
C3   第三支路；
CA、CB 电气线路。
具体实施方式
以下，结合附图，对本发明的具体实施方式进行详细说明。以下实施例将有助于本领域的技术人员进一步理解本发明，但不旨在对本发明的保护范围加以非必要的限制。
本发明的电动飞机综合电推进系统100独立于挂接多个电气化负载的电气化的机载系统即机载电能系统，其如图1所示是将集成在电动飞机10的机翼11上的分布式电推进系统130和电传飞控作动系统140综合成与统一的供电源110连接，并通过能量管理部120为电动飞机的飞行提供统一且能相互协调的电力能源供给，同时能够有效降低电动飞机的重量和能量能耗。
所述能量管理部120由后面描述的多个配电盘箱121对统一的供电源110的电能进行管理、分配和传输，从而向分布式电推进系统130和电传飞控作动系统140供电。所述分布式电推进系统130是以后面描述的多套电推进单元131为推力产生装置，为飞机飞行提供推力。所述电传飞控作动系统140是以后面描述的多套电传飞控作动单元141改变副翼、襟翼等的倾角，为飞机提供不同飞行姿态。
图2示出了布置在电动飞机10上的本发明的电动飞机综合电推进系统100的物理结构，图3是布置在电动飞机10上的本发明的电动飞机综合电推进系统100的系统架构图。
更具体来说，如图2所示，本发明的电动飞机综合电推进系统100呈左、右对称地安装在电动飞机10的左、右两侧的机翼11上，并且在每一侧的机翼11上各自包括至少一个(图2中示出为一个)配电盘箱121、构成分布式电推进系统130的多套(图2中示出为四套)电推进单元131、构成电传飞控作动系统140的多套(图2中示出为 四套)电传飞控作动单元141，如图3所示，位于每一侧的机翼11上的多套电推进单元131以及多套电传飞控作动单元141分别与位于该侧的配电盘箱121连接。
多套电推进单元131呈左、右对称地配置在机翼11的前侧边沿，并由各自配电盘箱121中的第一(例如270V高压)直流汇流条121H供电。每套电推进单元131包括高比功率三相逆变器131a、高比功率电机131b和推进装置131c。高比功率三相逆变器131a控制电推进单元131输出的推力。高比功率电机131b将三相电能转变为机械转速和转矩。推进装置131c采用螺旋桨或涵道风扇，将电机输出的转速和转矩转化为推力。
此外，多套电传飞控作动单元141呈左、右对称地配置在机翼11的后侧边沿，并由各自配电盘箱121中的第二(例如28V低压)直流汇流条121L供电。每套电传飞控作动单元141包括三相逆变器141a、伺服电机141b和作动筒141c。伺服电机141b根据三相逆变器141a的控制命令，输出一定转速。控制作动筒141c偏转副翼、襟翼等的舵面，从而调整飞行姿态。
在此，需要说明的是，在本发明所图示的实施例中，电动飞机10总体架构较小，因此，电传飞控作动单元141例如由28V低压直流汇流条供电，而当电动飞机10较大时，电传飞控作动单元141的供电汇流条电压等级将上升，例如采用270V高压直流汇流条供电，这取决于电动飞机10的具体大小和电传飞控作动单元141的功率。
另外，在每一侧的机翼11上还各自包括多套(图2中示出为两套)第一电池系统111和多套(图2中示出为两套)第二电池系统112，如图3所示，位于每一侧的机翼11上的多套第一电池系统111和多套第二电池系统112分别与配电盘箱121连接。位于左、右两侧的机翼11上的这些第一电池系统111和这些第二电池系统112以左、右对称的方式安装来构成图1所示的统一的供电源110，用于对左、右两侧的配电盘箱121进行供电。
如图3所示，第一电池系统111例如是高比能量锂电池系统，其具有锂电池包111a和电源变换设备111b。
虽未图示，锂电池包111a由若干锂电池单元和两套独立的电池管理系统、热管理系统、外壳、继电器、传感器等组成。每个锂电池单元由一个及以上锂电池单体组成。两套电池管理系统分别通过不同的方法监控锂电池单元的状态，并具备通信和控制功能。热管理系统能够动态调整锂电池包111a的温度，保证锂电池包111a的工作温度在合适的范围内。锂电池外壳具有防火、防爆等功能。继电器可以开通和切断锂电池单元的高压电气回路，实现对外的通断。
电源变换设备111b是DC/DC变换器，可以根据锂电池包的状态和飞行功率的需求进行充放电控制。
另外，如图3所示，第二电池系统112例如是应急锂电池系统，其具有锂电池包112a和能充放电的DC/DC变换器112b，用于应急情况下综合电推进系统关键设备的供电。在使用第一电池系统111作为电动飞机综合电推进系统100的供电源进行供电时，能够通过将DC/DC变换器112b切换为充电模式来对锂电池包112a进行充电，并且在第二电池系统112接近满电时进行涓流充电。另一方面，在使用第二电池系统112作为电动飞机综合电推进系统100的供电源进行供电时，能够将DC/DC变换器112b切换为放电模式来将锂电池包112a的电力输出。
另外，电动飞机综合电推进系统100如图3所示通过使位于左、右两侧的机翼11上的配电盘箱121各自的至少一个第一直流汇流条121H和至少一个第二直流汇流条121L分别通过电气线路CA、CB连接，并且位于左、右两侧的机翼11上的配电盘箱121构成图1所示的对供电源110(多套第一电池系统111和多套第二电池系统112)的电能进行管理、分配和传输的能量管理部120，由此能利用统一的供电源110通过相应一侧的配电盘箱121的控制对对应的各套电推进单元131和各套电传飞控作动单元141的电源供给，由此提供了多路冗 余备份的供电通道，提高了供电源110的容错性能。
另外，位于左、右两侧的机翼11上的电推进单元131构成图1所示的分布式电推进系统130。并且，位于左、右两侧的机翼11上的电传飞控作动单元141构成图1所示的电传飞控作动系统140。
图4是本发明的电动飞机综合电推进系统10的、位于一侧的机翼11的系统架构图，其中，示出了配电盘箱121内部的具体构成。
更具体来说，如图4所示，每个配电盘箱121包括多个(图4中为两个)串联连接的第一直流汇流条121H和多个(图4中为两个)串联连接的第二直流汇流条121L，多个串联连接的第一直流汇流条121H中，两两电气相邻的第一直流汇流条121H之间分别通过第一支路C1相互连接，多个串联连接的第二直流汇流条121L中，两两电气相邻的第二直流汇流条121L之间分别通过第二支路C2相互连接，另外，多个第一直流汇流条121H与多个第二直流汇流条121L逐个对应地通过设置有电压转换设备121a的第三支路C3相互连接。所述电压转换设备121a例如是DC/DC变换器，将第一直流汇流条121H所提供电压的直流电转换为第二直流汇流条121L所需电压的直流电。
另外，多套第一电池系统111与每个配电盘箱121中的多个第一直流汇流条121H分别连接，并且每个配电盘箱121中的每个第一直流汇流条121H均与多套(图4中示出为两套)电推进单元131连接。此外，多套第二电池系统112与每个配电盘箱121中的多个第二直流汇流条121L分别连接，并且每个配电盘箱121中的每个第二直流汇流条121L均与多套(图4中示出为两套)电传飞控作动单元141连接。
另外，虽未图示，每个配电盘箱121还包括汇流条接触器、控制继电器、热断路器等开关设备。另外，每条支路(第一支路C1、第二支路C2以及第三支路C3)均具有接触器或断路器等线路开关保护设备。配电盘箱121内部采用冗余容错的电网结构，用于保证电动飞机综合电推进系统100电网的安全可靠运行。
在每个配电盘箱121中，当正常供电时，通过每套第一电池系统111向配电盘箱121中的对应的第一直流汇流条121H供电，以确保与各第一直流汇流条121H连接的电推进单元131的运行，接着电流会经由对应的第三支路C3流入与各所述第一直流汇流条电气相邻连接的第二直流汇流条121L，以确保与各第二直流汇流条121L连接的电传飞控作动单元141的运行。此时的供电路径即第一供电路径是 即，在正常供电的情况下，每个第一直流汇流条121H由每套第一电池系统111直接供电，而每个第二直流汇流条121L从与其电气相邻连接的各第一直流汇流条121H经由电压转换设备121a供电。
另外，在每个配电盘箱121中，多个第一直流汇流条121H具有冗余供电的能力，当多套第一电池系统111的正常供电支路存在部分失效时，为确保与失效的第一电池系统111对应的第一直流汇流条121H能够正常工作，通过邻近的未失效的第一电池系统111向其供电，即，邻近的未失效的第一电池系统111首先向配电盘箱121中的与未失效的第一电池系统111对应的第一直流汇流条121H供电，以确保与该第一直流汇流条121H对应的电推进单元131的运行，接着电流会经由对应的第三支路C3流入与该第一直流汇流条121H电气相邻连接的第二直流汇流条121L，以确保与该第二直流汇流条121L连接的电传飞控作动单元141的运行；与此同时，电流会从与未失效的第一电池系统111对应的第一直流汇流条121H经由对应的第一支路C1向与失效的第一电池系统111对应的第一直流汇流条121H(又称“作为被冗余供电对象的第一直流汇流条”)供电，以确保与该第一直流汇流条121H(即，作为被冗余供电对象的第一直流汇流条121H)连接的电推进单元131的运行，接着，若是作为被冗余供电对象的第一直流汇流条121H与电气相邻连接的第二直流汇流条121L(又称“作为供电对象的第二直流汇流条”)之间的第三支路C3为通路(断路器闭 合)，则电流会从作为被冗余供电对象的第一直流汇流条121H经由对应的第三支路C3流入作为供电对象的第二直流汇流条121L，相反，若是作为被冗余供电对象的第一直流汇流条121H与作为供电对象的第二直流汇流条121L之间的第三支路C3为断路(断路器打开)，则电流无法从相应的第三支路C3流至作为供电对象的第二直流汇流条121L，此时，经由对应的第二支路C2从与作为供电对象的第二直流汇流条121L电气相邻连接的第二直流汇流条121L(该第二直流汇流条121L与对应的第一直流汇流条121H之间的对应的第三支路C3为通路(断路器闭合))供电。即，在正常供电支路存在部分失效的情况下，切换为与第一供电路径不同的第二供电路径，无法从失效的第一电池系统111获得电力的第一直流汇流条121H可以经由其他的能从未失效的第一电池系统111正常获得电力的第一直流汇流条121H间接地获得电力，而第二直流汇流条121L则可以根据与电气相邻连接的第一直流汇流条121H之间的第三支路C3的通断，经由与其电气相邻连接的第一直流汇流条121H或其他的第二直流汇流条121L供电。
在多套第一电池系统111的正常供电支路全部失效的情况下，此时为应急情况，向第一直流汇流条121H的供电不能得到保证，但为了避免电动飞机10完全失控，通过第二电池系统112向对应的第二直流汇流条121L供电，以确保与各第二直流汇流条121L连接的电传飞控作动单元141的运行。此时的供电支路即应急供电支路是 另外，在每个配电盘箱121中，多个第二直流汇流条121L同样具有冗余供电的能力，当多套第二电池系统112的应急供电支路存在部分失效时，为确保对应的第二直流汇流条121H能够正常工作，通过邻近的未失效的第二电池系统112向其供电，即，通过邻近的未失效的第二电池系统112首先向配电盘箱121中的与其对应的第二直流汇流条121L供电；与此同时电流会从与未失效的第二电池系统112对应的第二直流汇流条121L经由对应的第二支路C2向与失效的第二电池系统112对应的第二直流汇流条 121L(又称“作为被冗余供电对象的第二直流汇流条”)供电，以最大限度确保所述电传飞控作动系统140的运行。即，在正常供电支路全部失效的情况下，切换为与第一供电路径和第二供电路径不同的第三供电路径或第四供电路径，第二直流汇流条121L可以经由对应的第二电池系统112或是与其电气相邻连接的其他的第二直流汇流条121L供电。
换言之，本发明的电动飞机综合电推进系统100的每个配电盘箱121内部具有多个以第一电池系统111为供电源的通电通道，在正常供电的情况下，多个第一直流汇流条121H的每一个分别由各自直接连接的第一电池系统111供电，多个第二直流汇流条121L的每一个也分别经由直接连接的电压转换设备121a从电气相邻连接的第一直流汇流条121H获得电力，当某个供电通道中出现故障时，与该通电通道中的第一电池系统111对应的第一直流汇流条121H会切换至由邻近的通电通道供电。当发生应急情况时，多个第二电池系统112为多个第二直流汇流条121L上的电传飞控作动单元141提供应急电源，并且多个第二直流汇流条121L中的每一个具体冗余供电的能力，能由多个第二电池系统112供电。
再换一种说法，多个第一直流汇流条121H中的每一个第一直流汇流条121H(又称“目标第一直流汇流条”)允许由与该目标第一直流汇流条直接连接的所述第一电池系统111和与该第一直流汇流条121H(即，目标第一直流汇流条)电气相邻连接的第一直流汇流条选择一条供电路径。多个第二直流汇流条121L中的每一个第二直流汇流条121L(又称“目标第二直流汇流条”)允许由与该目标第二直流汇流条直接连接的所述第二电池系统112、与该第二直流汇流条121(即，目标第二直流汇流条)电气相邻连接的第一直流汇流条以及与该第二直流汇流条121(即，目标第二直流汇流条)电气相邻连接的第二直流汇流条选择一条供电路径。
需要说明的是，在与目标第一直流汇流条或目标第二直流汇流 条电气相邻的第一直流汇流条和/或第二直流汇流条亦无法从第一电池系统111或第二电池系统112直接连接(供电)的情况下，则再次选择下一个电气相邻连接的第一直流汇流条和/或第二直流汇流条。
熟悉本领域的技术人员易于想到其它的优点和修改。因此，在其更宽泛的上来说，本实用新型并不局限于这里所示和所描述的具体细节和代表性实施例。因此，可以在不脱离如所附权利要求书及其等价物所限定的总体发明概念的精神或范围的前提下做出修改。
在本发明的电动飞机综合电推进系统100中，配电盘箱121、第一电池系统111、第二电池系统112、电推进单元131、电传飞控作动单元141、第一直流汇流条121H、第二直流汇流条121L等的数量不局限于图2中所示数量。
例如，在本发明的电动飞机综合电推进系统100中，不局限于在左、右两侧的机翼11上各自仅配设一个配电盘箱121，也可以各自配设两个以上的配电盘箱121。另外，在左、右两侧的机翼11上配设的配电盘箱121的数量优选的是相同的，但也可以是不同的，例如可以在一侧的机翼11上配设一个配电盘箱121，而在另一侧的机翼11上配设两个配电盘箱121。
另外，对于一个配电盘箱121，只要具有两个以上的第一电池系统111和两套以上的第二电池系统112，并且每个第一电池系统111对应有两套以上的电推进单元131，而每套第二电池系统112对应有两套以上的电传飞控作动单元141即可。
另外，在单侧的机翼11上配设有多个配电盘箱121的情况下，只要其中至少一个配电盘箱121具备前述结构即可，并不必然要求每个配电盘箱121均具备前述结构。
另外，在图4所示的配电盘箱121的结构中，具有两个第一直流汇流条121H和两个第二直流汇流条121L，他们配置成：一个(例如图4中的左上方的)第一直流汇流条121H与电气相邻的(例如图4 中的右上方的)另一个第一直流汇流条121H通过第一支路C1连接；一个(例如图4中的左上方的)第一直流汇流条121H与电气相邻的(例如图4中的左下方的)一个第二直流汇流条121L通过第三支路C3连接；一个(例如图4中的左下方的)第二直流汇流条121L与电气相邻的(例如图4中的右下方的)另一个第二直流汇流条121L通过第二支路C2连接；另一个(例如图4中的右上方的)第一直流汇流条121H与电气相邻的(例如图4中的右下方的)第二直流汇流条121L通过第三支路C3连接，但本发明的电动飞机综合电推进系统100的配电盘箱121的结构中，也可以具有三个以上第一直流汇流条121H和三个以上第二直流汇流条121L，以三个的情况为例，他们可以配置成：一个(例如左上方的)第一直流汇流条121H与电气相邻的(例如中间上方的)另一个第一直流汇流条121H通过第一支路C1连接，该另一个第一直流汇流条121H与电气相邻的(例如右上方的)又一个第一直流汇流条121H也通过第一支路C1连接；一个(例如左下方的)第二直流汇流条121L与电气相邻的(例如中间下方的)另一个第二直流汇流条121L通过第二支路C2连接，该另一个第二直流汇流条121L与电气相邻的(例如右下方的)又一个第二直流汇流条121L也通过第二支路C2连接；一个(例如左上方的)第一直流汇流条121H与电气相邻的一个(例如左下方的)第二直流汇流条121L通过第三支路C3连接，另一个(例如中间上方的)第一直流汇流条121H与电气相邻的另一个(例如中间下方的)第二直流汇流条121L通过第三支路C3连接，又一个(例如右上方的)第一直流汇流条121H与电气相邻的又一个(例如右下方的)第二直流汇流条121L通过第三支路C3连接。
此处，在本申请的说明书中，“与特定的部件B电气相邻连接的特定的部件A”、“特定的部件A与(和特定的部件A)电气相邻的特定的部件B连接”等表述指的是在特定的部件A与特定的部件B之间的支路中不存在其他的部件A或部件B，例如，在图4中，左上方的第一直流汇流条121H与左下方的第二直流汇流条121L以及与右上 方的第一直流汇流条121H是电气相邻连接的，但与右下方的第二直流汇流条121L并不是电气相邻连接的。此外，在特定的部件A与特定的部件B之间的支路中可以具有部件A与部件B之外的部件C，例如，在图4中，电气相邻连接的左上方的第一直流汇流条121H与左下方的第二直流汇流条121L之间可以具有电压转换设备121a(以及断路器)等。另外，此时电气相邻连接的部件A与部件B并不必然是指物理上彼此相邻或是靠得很近的两个部件，也可以是物理上相隔较远的两个部件，只要满足上面的情形。
另外，在图3所示的电动飞机综合电推进系统的系统架构图，为避免线路的复杂化，位于左侧(例如图3中的上侧)的机翼11上的配电盘箱121与位于右侧(例如图3中的下侧)的机翼11上的配电盘箱121以仅使其中各自一个第一直流汇流条121H通过电气线路CA连接并且仅使其中各自一个第二直流汇流条121L通过电气线路CB连接的方式相互电气连接，但本发明不局限于此，也可以配置成使其中每个第一直流汇流条121H对应地通过电气线路CA连接并且使其中每个第二直流汇流条121L对应地通过另一电气线路CB连接。
另外，在本发明中，以第一电池系统111是高比能量锂电池系统，第二电池系统112是应急锂电池系统为例进行了说明，但本发明不局限于此，第一电池系统111和第二电池系统112可以是任意适合用于电动飞机的电池系统。此外，考虑到第一电池系统111作为主电源，第二电池系统112作为应急(备用)电源，优选的是，第一电池系统111的电容量大于第二电池系统112的电容量。同时，虽然在说明书的实施例中，列举了第二电池系统112是可充放电的电池系统，但第二电池系统112并不是必须具备充电功能，也可以是仅放电的电池系统。

Claims (13)

  1. 一种电动飞机综合电推进系统，其特征在于，
    所述电动飞机综合电推进系统独立于挂接多个电气化负载的机载电能系统，
    所述电动飞机综合电推进系统将集成在电动飞机的机翼上的分布式电推进系统和电传飞控作动系统综合成与统一的供电源连接，
    在正常供电时，以第一供电路径向所述分布式电推进系统和所述电传飞控作动系统供电，并且
    在所述正常供电发生部分失效的故障情况时或是发生全部失效的应急情况时，通过能量管理部，将向所述分布式电推进系统和所述电传飞控作动系统供电的路径切换成与所述第一供电路径不同的供电路径。
  2. 如权利要求1所述的电动飞机综合电推进系统，其特征在于，
    所述能量管理部是多个配电盘箱，
    多个所述配电盘箱中的至少一个所述配电盘箱包括多个串联连接的第一直流汇流条和多个串联连接的第二直流汇流条，其中：
    多个串联连接的所述第一直流汇流条中，两两电气相邻的所述第一直流汇流条之间分别通过第一支路相互连接；
    多个串联连接的所述第二直流汇流条中，两两电气相邻的所述第二直流汇流条之间分别通过第二支路相互连接；
    多个所述第一直流汇流条与多个所述第二直流汇流条逐个对应地通过设置有电压转换设备的第三支路相互连接，
    多个配电盘箱通过使各自的至少一个所述第一直流汇流条对应地通过电气线路相互连接，并且使各自的至少一个所述第二直流汇流条对应地通过另一电气线路相互连接，由此构成所述能量管理部。
  3. 如权利要求2所述的电动飞机综合电推进系统，其特征在于，
    所述统一的供电源包括多套第一电池系统和多套第二电池系统，
    每套所述第一电池系统对应地连接于每个第一直流汇流条，
    每套所述第二电池系统对应地连接于每个第二直流汇流条。
  4. 如权利要求3所述的电动飞机综合电推进系统，其特征在于，
    在向各个所述第一直流汇流条正常供电时，向所述分布式电推进系统和所述电传飞控作动系统供电的第一供电路径是：
    通过每套所述第一电池系统向所述配电盘箱中的对应的所述第一直流汇流条供电，以确保与各所述第一直流汇流条连接的构成所述分布式电推进系统的电推进单元的运行，接着电流会经由对应的所述第三支路流入与各所述第一直流汇流条电气相邻连接的各所述第二直流汇流条，以确保与各所述第二直流汇流条连接的构成所述电传飞控作动系统的电传飞控作动单元的运行。
  5. 如权利要求3所述的电动飞机综合电推进系统，其特征在于，
    在向各个所述第一直流汇流条正常供电发生部分失效时，即在多套所述第一电池系统的正常供电支路存在部分失效的故障情况下，向所述分布式电推进系统和所述电传飞控作动系统供电的第二供电路径是：
    通过邻近的未失效的第一电池系统向配电盘箱中的与其对应的第一直流汇流条供电，以确保与该第一直流汇流条连接的构成所述分布式电推进系统的电推进单元的运行，接着电流会经由对应的所述第三支路流入与该第一直流汇流条电气相邻连接的第二直流汇流条，以确保与该第二直流汇流条连接的构成所述电传飞控作动系统的电传飞控作动单元的运行，
    同时电流会从与未失效的第一电池系统对应的第一直流汇流条经由所述第一支路向与失效的第一电池系统对应的第一直流汇流条即作为被冗余供电对象的第一直流汇流条供电，以确保与作为所述被冗余供电对象的第一直流汇流条连接的电推进单元的运行，接着电流会从作为所述被冗余供电对象的第一直流汇流条经由对应的第三支路流入与该第一直流汇流条电气相邻连接的第二直流汇流条，或是从电气相邻连接的所述第二直流汇 流条经由对应的第二支路流入与作为所述被冗余供电对象的第一直流汇流条电气相邻连接的第二直流汇流条。
  6. 如权利要求3所述的电动飞机综合电推进系统，其特征在于，
    在向各个所述第一直流汇流条正常供电全部失效时，即在多套所述第一电池系统的正常供电支路全部失效的应急情况下，向所述电传飞控作动系统供电的第三供电路径是：
    通过所述第二电池系统向对应的所述第二直流汇流条供电，以确保与各第二直流汇流条连接的构成所述电传飞控作动系统的电传飞控作动单元的运行。
  7. 如权利要求6所述的电动飞机综合电推进系统，其特征在于，
    当多套所述第二电池系统的应急供电支路存在部分失效时，向所述电传飞控作动系统供电的第四供电路径是：
    通过邻近的未失效的第二电池系统向配电盘箱中的与其对应的第二直流汇流条供电，
    同时电流会从与未失效的第二电池系统对应的所述第二直流汇流条经由对应的所述第二支路向与失效的第二电池系统对应的第二直流汇流条即作为被冗余供电对象的第二直流汇流条供电。
  8. 如权利要求3所述的电动飞机综合电推进系统，其特征在于，
    多个所述第一直流汇流条中的每一个第一直流汇流条允许由所述第一电池系统和与该第一直流汇流条电气相邻连接的第一直流汇流条选择一条供电路径。
  9. 如权利要求8所述的电动飞机综合电推进系统，其特征在于，
    多个所述第二直流汇流条中的每一个第二直流汇流条允许由所述第二电池系统、与该第二直流汇流条电气相邻连接的第一直流汇流条以及与该第二直 流汇流条电气相邻连接的第二直流汇流条选择一条供电路径。
  10. 如权利要求3所述的电动飞机综合电推进系统，其特征在于，
    所述电动飞机综合电推进系统的各个配电盘箱内部具有多个以所述第一电池系统为主电源的通电通道，
    在正常供电的情况下，多个所述第一直流汇流条中的每一个分别由各自直接连接的所述第一电池系统供电，多个第二直流汇流条中的每一个也分别经由直接连接的所述电压转换设备从电气相邻连接的所述第一直流汇流条获得电力，
    当某个供电通道中出现故障的故障情况时，与该通电通道中的所述第一电池系统对应的所述第一直流汇流条切换至由邻近的通电通道供电，
    当发生应急情况时，多个所述第二电池系统作为备用电源为多个所述第二直流汇流条上的电传飞控作动单元提供应急电源，并且多个第二直流汇流条中的每一个具体冗余供电的能力，能由多个第二电池系统供电。
  11. 如权利要求3至10中任一项所述的电动飞机综合电推进系统，其特征在于，
    所述电动飞机综合电推进系统呈左、右对称地安装在所述电动飞机的左、右两侧的所述机翼上，
    在每一侧的所述机翼上各自包括一个配电盘箱、构成所述分布式电推进系统的四套电推进单元、构成所述电传飞控作动系统的四套电传飞控作动单元以及作为统一的供电源的两套第一电池系统和两套第二电池系统。
  12. 如权利要求11所述的电动飞机综合电推进系统，其特征在于，
    所述第一电池系统是高比能量锂电池系统，
    所述第二电池系统是应急锂电池系统。
  13. 如权利要求12所述的电动飞机综合电推进系统，其特征在于，
    所述应急锂电池系统具有能充放电的DC/DC变换器。
PCT/CN2023/085041 2022-09-22 2023-03-30 电动飞机综合电推进系统 WO2024060567A1 (zh)

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