CN115432190B - Comprehensive electric propulsion system of electric aircraft - Google Patents

Abstract

The comprehensive electric propulsion system for electric aircraft can integrate the distributed electric propulsion system and fly-by-wire control actuating system integrated on the wing, provide reliable, efficient and stable flying power for the electric aircraft, and reduce the weight and energy loss of the system. The electric aircraft comprehensive electric propulsion system is independent of an airborne electric energy system with a plurality of electrified loads, the electric aircraft comprehensive electric propulsion system integrates a distributed electric propulsion system and an electric flight control actuating system which are integrated on a wing of an electric aircraft into a unified power supply source, a first power supply circuit is used for supplying power to the distributed electric propulsion system and the electric flight control actuating system in normal power supply, and when a fault condition of partial failure of the normal power supply or an emergency condition of total failure occurs, the energy management part is used for switching a power supply path of the distributed electric propulsion system and the electric flight control actuating system into a power supply path different from the first power supply path.

Description

Comprehensive electric propulsion system of electric aircraft

Technical Field

The invention relates to an electric propulsion technology of an aircraft, in particular to an electric aircraft comprehensive electric propulsion system.

Background

With the development of electrification technology, the electric aircraft adopts a battery and a propulsion motor to replace the traditional fuel oil and engine driving mode of the aircraft, so that the noise and emission are reduced, and the energy utilization rate is improved. At present, the electric propulsion system is limited by motor technology and power distribution technology, and a plurality of electric propulsion units are arranged at the edge of an airplane wing and integrated into a distributed electric propulsion system, so that effective thrust is provided for an electric airplane under the condition of meeting the existing mature power electronic technology.

However, the moment generated by the thrust force varies from wing to wing, and if only a distributed electric propulsion system is used, the stability of the aircraft may be greatly and negatively affected.

Meanwhile, the specific energy of the battery limits the output capacity of the distributed electric propulsion system, and the aircraft must comprehensively consider the influence of thrust output and flight attitude on the flight performance, so that the limited electric power and energy resources are fully utilized, and the flight capacity of the aircraft is improved.

In addition, electrified flight control systems (also known as fly-by-wire actuation systems) have short-time, high-power charge-discharge characteristics, which result in much higher power transmission, distribution device weight and losses in the upstream branches than conventional flight control systems.

Disclosure of Invention

Therefore, the present invention has been made to solve the above-mentioned problems occurring in the prior art, and it is an object of the present invention to provide an electric aircraft integrated electric propulsion system, which integrates a distributed electric propulsion system and a fly-by-wire actuation system integrated on a wing, to provide a reliable, efficient and stable flying power for an electric aircraft, while reducing the weight and energy consumption of the system.

In order to achieve the above object, the present invention provides an electric aircraft integrated electric propulsion system, which is independent of an on-board electric power system for hooking a plurality of electrified loads, and is characterized in that the electric aircraft integrated electric propulsion system integrates a distributed electric propulsion system integrated on a wing of an electric aircraft and a fly-by-wire actuation system into a unified power supply source, and in normal power supply, a first power supply circuit is used for supplying power to the distributed electric propulsion system and the fly-by-wire actuation system, and in case of a failure of a partial failure of the normal power supply or an emergency of a total failure, a power supply path for supplying power to the distributed electric propulsion system and the fly-by-wire actuation system is switched to a different power supply path from the first path by an energy management part.

According to the above constitution, compared with the architecture of the prior art, the electric aircraft comprehensive electric propulsion system of the invention has the following beneficial effects:

(1) The electric aircraft comprehensive electric propulsion system and the airborne electric energy system are mutually independent, and the electric aircraft comprehensive electric propulsion system is simple to control, low in implementation difficulty and easy to implement and modify;

(2) The electric aircraft comprehensive electric propulsion system has the characteristics of a distributed energy system, reduces cables and conversion devices for equipment power supply, reduces the weight of the system, reduces the loss of electric energy and improves the reliability of the system;

(3) The electric aircraft comprehensive electric propulsion system integrates two system functions of thrust output (distributed electric propulsion system) and attitude control (fly-by-wire control actuating system), provides a basis for comprehensive and cooperative control of the two, and is beneficial to improving the performance, stability and energy utilization rate of the electric aircraft.

(4) The integrated electric propulsion system of the electric aircraft is connected with the distributed electric propulsion system and the fly-by-wire control actuation system by adopting a unified power supply, and when a part of main power supply fails, the power supply path which is different from a first power supply path in normal power supply can be utilized to supply power to the whole distributed electric propulsion system and the whole fly-by-wire control actuation system by utilizing the main power supply which is not failed, so that the functions of the two systems of thrust output and attitude control of the electric aircraft are guaranteed to the greatest extent. Even if all main power supplies fail (fail), even a part of standby power supplies fail (fail), the whole fly-by-wire control actuating system can be powered by the standby power supplies which are not failed, so that safe landing of the electric aircraft can be ensured through gesture control.

Preferably, the energy management unit is a plurality of distribution board boxes. The plurality of distribution board boxes can be wing installed on the left side and the right side symmetrically, or the number of distribution board boxes installed on one side of wing is different from the number of distribution board boxes installed on one side of wing.

In one example, for example, one panel box is included on each of the wings on each side. And, preferably, each of the distribution board boxes includes a plurality of first dc bus bars connected in series and a plurality of second dc bus bars connected in series. In the switchboard box with the structure, a plurality of first direct current bus bars connected in series are connected with each other by a first branch circuit, wherein every two adjacent first direct current bus bars are connected with each other by a first branch circuit; among the plurality of second direct current bus bars connected in series, every two of the second direct current bus bars which are electrically adjacent are connected with each other through a second branch; the plurality of first dc bus bars and the plurality of second dc bus bars are connected to each other by a third branch provided with a voltage conversion device, one by one, correspondingly. Among the plurality of distribution boards, one or more first direct current bus bars of one distribution board box are connected to one another by an electric line, respectively, and one or more second direct current bus bars of one distribution board box are connected to one another by an electric line, respectively, thereby constituting the energy management section.

Because the multiple distribution board boxes can be provided with multiple redundant backup power supply channels, the fault tolerance of the power supply source can be improved.

In addition to this one example, it is also possible to include more than one panel box on one side of the wing and more than two panel boxes on the other side of the wing. In the three or more distribution boxes on both sides, at least one of the distribution boxes may have the structure of the distribution box of the above-described one example, and the first dc bus bars may be connected to each other by an electric line between the three or more distribution boxes and the second dc bus bars may be connected to each other by another electric line, thereby constituting the energy management section.

Further preferably, the first dc bus bar in each of the switchboard boxes is connected to the first battery system, and the second dc bus bar in each of the switchboard boxes is connected to the second battery system, and the first battery system and the second battery system constitute a unified power supply.

In other words, each of the plurality of first dc bus bars allows one power supply path to be selected by the first battery system and the first dc bus bar that is electrically adjacent to the first dc bus bar. And, each of the plurality of second dc bus bars allows one power supply path to be selected by the second battery system, the first dc bus bar electrically connected adjacent to the second dc bus bar, and the second dc bus bar electrically connected adjacent to the second dc bus bar.

According to the above configuration, each of the switchboard boxes of the electric aircraft integrated electric propulsion system has a plurality of energizing passages using the first battery system as a main power source, and when an emergency occurs, the plurality of second battery systems serve as standby power sources to supply emergency power to the fly-by-wire operation units on the plurality of second dc bus bars.

In the case of normal power supply, each of the plurality of first direct current bus bars is supplied with power by the respective directly connected first battery system, and each of the plurality of second direct current bus bars also obtains power from the electrically adjacent connected first direct current bus bars via the directly connected voltage conversion device, respectively. Thus, when normally powering the respective first dc bus bars, the first power supply path for powering the distributed electric propulsion system and the fly-by-wire actuation system is: power is supplied to corresponding first dc bus bars in the switchboard box by each set of first battery systems to ensure operation of electric propulsion units connected to the first dc bus bars, which constitute a distributed electric propulsion system, and then current flows to the second dc bus bars connected to the first dc bus bars electrically adjacent to each other via corresponding third branches to ensure operation of electric flight control actuation units connected to the second dc bus bars, which constitute an electric flight control actuation system.

In addition, when a fault condition occurs in which a fault in a certain power supply channel, the first direct current bus bar corresponding to the first battery system in the power supply channel is switched to be supplied with power from an adjacent power supply channel. Therefore, when the normal power supply to each first direct current bus bar is partially failed, namely, when the normal power supply branch of the plurality of sets of first battery systems has a fault that is partially failed, the second power supply path for supplying power to the distributed electric propulsion system and the fly-by-wire actuation system is as follows: the operation of the electric propulsion units forming the distributed electric propulsion system connected with the first direct current bus bar is ensured by supplying power to the corresponding first direct current bus bar in the distribution panel box through the adjacent first battery system which does not fail, then the current flows into the second direct current bus bar which is connected with the first direct current bus bar in an electric adjacent way through the corresponding third branch circuit, so as to ensure the operation of the electric fly control operation unit forming the electric fly control operation system connected with the second direct current bus bar, and simultaneously the current is supplied from the first direct current bus bar corresponding to the first battery system which does not fail to the first direct current bus bar corresponding to the first battery system which does not fail through the first branch circuit, so as to ensure the operation of the electric propulsion units connected with the first direct current bus bar which is a redundant power supply object to flow into the second direct current bus bar which is connected with the first direct current bus bar in an electric adjacent way through the corresponding third branch circuit, or the current flows into the first direct current bus bar which is a redundant power supply object through the corresponding second direct current bus bar which is connected with the second direct current bus bar in an electric adjacent way through the corresponding second branch circuit.

That is, as long as a certain first dc bus is not completely electrically disconnected from the first battery system and from the first dc bus electrically connected to the first dc bus, the operation of the electric propulsion unit connected to the first dc bus can be ensured, whereby both the thrust output and the attitude control system functions of the electric aircraft can be ensured to the maximum extent.

When all the normal power supply to each first direct current bus bar fails, namely, in the emergency situation that all the normal power supply branches of the plurality of sets of first battery systems fail, a third power supply path for supplying power to the fly-by-wire control actuation system is as follows: and supplying power to the corresponding second direct current bus bars through the second battery system so as to ensure the operation of the fly-by-wire actuation units connected with the second direct current bus bars and forming the fly-by-wire actuation system.

Preferably, each of the plurality of second dc bus bars is capable of being powered by a plurality of second battery systems with a specific redundant power capability. At this time, when the emergency power supply branch of the plurality of sets of second battery systems has partial failure, the fourth power supply path for supplying power to the fly-by-wire actuation system is: and supplying power to the second direct current bus bars corresponding to the non-failure second battery systems in the distribution board box through the adjacent non-failure second battery systems, and supplying current from the second direct current bus bars corresponding to the non-failure second battery systems to the second direct current bus bars corresponding to the failure second battery systems, namely the second direct current bus bars serving as redundant power supply objects, through the corresponding second branches, so that the operation of the fly-by-wire control actuation system is ensured to the greatest extent.

That is, as long as a certain second dc bus bar is not completely electrically disconnected from the second battery system, the first dc bus bar electrically connected to the second dc bus bar, and the second dc bus bar electrically connected to the second dc bus bar, the operation of the fly-by-wire operation unit connected to the second dc bus bar can be ensured. Therefore, the two system functions of attitude control of the electric aircraft can be guaranteed to the greatest extent, and the risk of out-of-control of the electric aircraft under emergency conditions is reduced.

In one embodiment of the invention, the electric aircraft comprehensive electric propulsion system is symmetrically arranged on the left and right wings of the electric aircraft in a left-right symmetrical way, and each wing on each side comprises a distribution board box, four sets of electric propulsion units forming a distributed electric propulsion system, four sets of electric flight control actuation units forming an electric flight control actuation system, and two sets of first battery systems and two sets of second battery systems serving as uniform power sources. However, it should be understood that this is only for a particular architecture of a particular type or type/size of electric aircraft, and that the specific number of the present invention should not be limited to this particular architecture in a broader sense.

The first battery system is, for example, a high specific energy lithium battery system. In addition, the second battery system is, for example, an emergency lithium battery system.

Meanwhile, the second battery system (emergency lithium battery system) may be a battery system having a charge/discharge function, or may be a battery system having only a discharge function. In the former case, it is preferable to have a DC/DC converter that can be charged and discharged.

Drawings

FIG. 1 is a schematic block diagram of an electric aircraft integrated electric propulsion system of the present invention.

Fig. 2 is a physical schematic of the electric aircraft integrated electric propulsion system of the present invention disposed on an electric aircraft.

Fig. 3 is a system architecture diagram of the electric aircraft integrated electric propulsion system of the present invention disposed on an electric aircraft.

Fig. 4 is a system configuration diagram of a wing on one side of the electric aircraft integrated electric propulsion system of the present invention, wherein a specific configuration of the interior of the switchboard box is shown.

(symbol description)

10. An electric aircraft;

11. a wing;

100. an electric aircraft integrated electric propulsion system;

110. a power supply;

111. a first battery system;

111a lithium battery pack;

111b power conversion apparatus;

112. a second battery system;

112a lithium battery pack;

112b DC/DC converter;

120. an energy management unit;

121. a distribution board box;

121a voltage conversion device;

121H first dc bus;

121L of a second dc bus bar;

130. a distributed electric propulsion system;

131. an electric propulsion unit;

131a high specific power three-phase inverter;

131b high specific power motor;

131c propulsion means;

140. fly-by-wire actuation system;

141. fly-by-wire actuation unit;

141a three-phase inverter;

141b servo motor;

141c rams;

c1 A first branch;

c2 A second branch;

c3 A third branch;

CA. CB electrical circuit.

Detailed Description

Hereinafter, embodiments of the present invention will be described in detail with reference to the accompanying drawings. The following examples will assist those skilled in the art in further understanding the present invention, but are not intended to unnecessarily limit the scope of the present invention.

The electric aircraft integrated electric propulsion system 100 according to the present invention is independent of an on-board system, i.e., an on-board electric power system, to which a plurality of electric loads are connected, and as shown in fig. 1, integrates a distributed electric propulsion system 130 and an electric flight control system 140 integrated on a wing 11 of an electric aircraft 10 to be connected to a unified power supply 110, and provides unified and mutually coordinated electric power energy supply for the flight of the electric aircraft through an energy management part 120, and at the same time, can effectively reduce the weight and energy consumption of the electric aircraft.

The energy management part 120 manages, distributes and transmits electric energy of the unified power supply 110 by a plurality of distribution board boxes 121 described later, thereby supplying power to the distributed electric propulsion system 130 and the fly-by-wire actuation system 140. The distributed electric propulsion system 130 uses a plurality of sets of electric propulsion units 131 as thrust generating devices, which are described later, to provide thrust for the aircraft to fly. The fly-by-wire actuation system 140 changes the inclination angle of ailerons, flaps, etc. with a plurality of fly-by-wire actuation units 141 described later to provide different flight attitudes for the aircraft.

Fig. 2 shows the physical structure of the electric aircraft integrated electric propulsion system 100 of the present invention arranged on the electric aircraft 10, and fig. 3 is a system configuration diagram of the electric aircraft integrated electric propulsion system 100 of the present invention arranged on the electric aircraft 10.

More specifically, as shown in fig. 2, the electric aircraft integrated electric propulsion system 100 of the present invention is symmetrically installed on the left and right wings 11 of the electric aircraft 10 in a left-right direction, and each of the wings 11 on each side includes at least one (one is shown in fig. 2) switchboard box 121, a plurality of sets (four is shown in fig. 2) of electric propulsion units 131 constituting the distributed electric propulsion system 130, a plurality of sets (four is shown in fig. 2) of electric flight control actuation units 141 constituting the electric flight control actuation system 140, and as shown in fig. 3, a plurality of sets of electric propulsion units 131 and a plurality of sets of electric flight control actuation units 141 on the wing 11 on each side are respectively connected to the switchboard boxes 121 on that side.

The plurality of electric propulsion units 131 are symmetrically disposed on the front side edge of the wing 11 in a left-right direction and are supplied with power from the first (e.g., 270V high voltage) dc bus bar 121H in the respective switchboard boxes 121. Each set of electric propulsion units 131 comprises a high specific power three-phase inverter 131a, a high specific power motor 131b and propulsion means 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 electric energy into mechanical rotational speed and torque. The propulsion device 131c converts the rotational speed and torque output from the motor into thrust by using a propeller or a ducted fan.

Further, the plurality of fly-by-wire actuation units 141 are symmetrically disposed on the rear edge of the wing 11 in left-right directions and are supplied with power from the second (e.g., 28V low voltage) dc bus bar 121L in the respective switchboard boxes 121. Each set of fly-by-wire actuation units 141 includes a three-phase inverter 141a, a servo motor 141b, and an actuator cylinder 141c. The servo motor 141b outputs a certain rotation speed according to a control command of the three-phase inverter 141 a. The control actuator 141c deflects the control surface of the aileron, flap, or the like, thereby adjusting the attitude of the flight.

It should be noted here that in the illustrated embodiment of the present invention, the overall architecture of the electric aircraft 10 is small, and therefore, the fly-by-wire actuation unit 141 is powered by, for example, a 28V low voltage dc bus, whereas when the electric aircraft 10 is large, the voltage level of the power bus of the fly-by-wire actuation unit 141 will rise, for example, using a 270V high voltage dc bus, depending on the specific size of the electric aircraft 10 and the power of the fly-by-wire actuation unit 141.

In addition, a plurality of sets (two sets are shown in fig. 2) of first battery systems 111 and a plurality of sets (two sets are shown in fig. 2) of second battery systems 112 are also respectively included on each side of the wing 11, and as shown in fig. 3, the plurality of sets of first battery systems 111 and the plurality of sets of second battery systems 112 on each side of the 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 and right symmetrical manner to constitute a unified power supply 110 shown in fig. 1 for supplying power to the right and left switchboard boxes 121.

As shown in fig. 3, the first battery system 111 is for example a high specific energy lithium battery system,

which has a lithium battery pack 111a and a power conversion device 111b.

Although not shown, the lithium battery pack 111a is composed of a plurality of lithium battery cells and two independent battery management systems, thermal management systems, housings, relays, sensors, and the like. Each lithium battery unit consists of one or more lithium battery cells. The two sets of battery management systems monitor the states of the lithium battery units through different methods respectively and have communication and control functions. The thermal management system can dynamically adjust the temperature of the lithium battery pack 111a, and ensure that the operating temperature of the lithium battery pack 111a is within a suitable range. The lithium battery shell has the functions of fire prevention, explosion prevention and the like. The relay can open and cut off a high-voltage electric loop of the lithium battery unit to realize external on-off.

The power conversion device 111b is a DC/DC converter, and can perform charge and discharge control according to the state of the lithium battery pack and the flying power demand.

In addition, as shown in fig. 3, the second battery system 112 is, for example, an emergency lithium battery system having a lithium battery pack 112a and a DC/DC converter 112b capable of charging and discharging for supplying power to key devices of the integrated electric propulsion system in an emergency situation. When the first battery system 111 is used as a power supply source for the electric aircraft integrated electric propulsion system 100, the lithium battery pack 112a can be charged by switching the DC/DC converter 112b to a charging mode, and trickle charging is performed when the second battery system 112 is near full power. On the other hand, when the second battery system 112 is used as a power supply source for the electric aircraft integrated electric propulsion system 100, the DC/DC converter 112b can be switched to the discharge mode to output electric power from the lithium battery pack 112 a.

In addition, in the electric aircraft integrated electric propulsion system 100, as shown in fig. 3, at least one first dc bus bar 121H and at least one second dc bus bar 121L of each of the switchboard boxes 121 on the left and right wings 11 are connected through electric lines CA and CB, respectively, and the switchboard boxes 121 on the left and right wings 11 form the energy management part 120 shown in fig. 1 for managing, distributing and transmitting electric energy of the power supply source 110 (the plurality of first battery systems 111 and the plurality of second battery systems 112), thereby enabling the power supply of the corresponding sets of electric propulsion units 131 and sets of electric flight control actuation units 141 through the control of the switchboard boxes 121 on the respective sides by using the unified power supply source 110, thereby providing a power supply channel for multiple redundancy backup and improving the fault tolerance performance of the power supply source 110.

In addition, the electric propulsion units 131 on the wings 11 on the left and right sides constitute the distributed electric propulsion system 130 shown in fig. 1. Further, fly-by-wire actuation units 141 on the wings 11 on the left and right sides constitute a fly-by-wire actuation system 140 shown in fig. 1.

Fig. 4 is a system configuration diagram of the wing 11 on one side of the electric aircraft integrated electric propulsion system 10 of the present invention, in which a specific configuration of the inside of the switchboard box 121 is shown.

More specifically, as shown in fig. 4, each switchboard box 121 includes a plurality of (two in fig. 4) first dc bus bars 121H connected in series and a plurality of (two in fig. 4) second dc bus bars 121L connected in series, among the plurality of first dc bus bars 121H connected in series, the first dc bus bars 121H electrically adjacent to each other in pairs are connected to each other by a first branch C1, among the plurality of second dc bus bars 121L connected in series, the second dc bus bars 121L electrically adjacent to each other in pairs are connected to each other by a second branch C2, and further, the plurality of first dc bus bars 121H and the plurality of second dc bus bars 121L are connected to each other by a third branch C3 provided with a voltage conversion device 121 a. The voltage conversion device 121a is, for example, a DC/DC converter that converts the direct current of the voltage supplied from the first direct current bus bar 121H into the direct current of the voltage required by the second direct current bus bar 121L.

In addition, the plurality of sets of first battery systems 111 are connected to the plurality of first dc bus bars 121H in each of the switchboard boxes 121, respectively, and each of the first dc bus bars 121H in each of the switchboard boxes 121 is connected to the plurality of sets (two sets are shown in fig. 4) of electric propulsion units 131. Further, the plurality of sets of second battery systems 112 are connected to the plurality of second dc bus bars 121L in each of the switchboard boxes 121, respectively, and each of the second dc bus bars 121L in each of the switchboard boxes 121 is connected to the plurality of sets (two sets are shown in fig. 4) of fly-by-wire actuation units 141.

Further, although not shown, each of the switchboard boxes 121 further includes a switching device such as a bus bar contactor, a control relay, a thermal breaker, and the like. In addition, each branch (the first branch C1, the second branch C2, and the third branch C3) has a line switch protection device such as a contactor or a circuit breaker. The inside of the distribution board box 121 adopts a redundant fault-tolerant power grid structure, so as to ensure the safe and reliable operation of the power grid of the electric aircraft comprehensive electric propulsion system 100.

In each switchboard 121, when power is normally supplied, the corresponding first dc bus bar 121H in the switchboard 121 is supplied by each set of first battery systems 111 to ensure the operation of the electric propulsion units 131 connected to the respective first dc bus bars 121H, and then current flows into the second dc bus bars 121L electrically adjacent to the respective first dc bus bars via the corresponding third branches C3 to ensure the operation of the fly-by-wire actuation units 141 connected to the respective second dc bus bars 121L. The first power supply path at this time is the first battery system 111

First direct-current bus bar 121H->

Third branch C3->

And a second dc bus bar 121L. I.e. in normal power supplyEach first dc bus bar 121H is directly supplied by each set of first battery systems 111, and each second dc bus bar 121L is supplied from each first dc bus bar 121H connected electrically adjacent thereto via the voltage conversion device 121 a.

In addition, in each switchboard box 121, the plurality of first dc bus bars 121H have the capability of redundant power supply, when there is a partial failure in the normal power supply branch of the plurality of first battery systems 111, in order to ensure that the first dc bus bar 121H corresponding to the failed first battery system 111 can work normally, power is supplied to it through the adjacent non-failed first battery system 111, that is, the adjacent non-failed first battery system 111 first supplies power to the first dc bus bar 121H corresponding to the non-failed first battery system 111 in the switchboard box 121 first to ensure the operation of the electric propulsion unit 131 corresponding to the first dc bus bar 121H, and then current flows into the second dc bus bar 121L electrically adjacent to the first dc bus bar 121H via the corresponding third branch C3 to ensure the operation of the fly-by-wire actuation unit 141 connected to the second dc bus bar 121L; at the same time, current is supplied from the first dc bus bar 121H corresponding to the first battery system 111 that has not failed to the first dc bus bar 121H corresponding to the first battery system 111 that has failed (also referred to as "first dc bus bar that is the object to be supplied with redundancy") via the corresponding first branch C1 to ensure the operation of the electric propulsion unit 131 connected to the first dc bus bar 121H (that is, the first dc bus bar 121H that is the object to be supplied with redundancy), then, if the third branch C3 between the first dc bus bar 121H that is the object to be supplied with redundancy and the second dc bus bar 121L that is electrically adjacent to the second dc bus bar 121L (also referred to as "second dc bus bar that is the object to be supplied with power") is a passage (circuit breaker is closed), current flows from the first dc bus bar 121H that is the object to be supplied with redundancy via the corresponding third branch C3 to the second dc bus bar that is the object to be supplied with redundancy), and if the third branch C3 between the first dc bus bar 121H that is the object to be supplied with redundancy and the second dc bus bar 121L that is the corresponding to be supplied with power is a circuit breaker is not able to open (the circuit breaker is the corresponding second dc bus bar) from the corresponding second dc bus bar 121L). That is, when there is a partial failure in the normal power supply branch, the first dc bus bar 121H, which is switched to a second power supply path different from the first power supply path, and which cannot obtain power from the failed first battery system 111, may indirectly obtain power via the other first dc bus bar 121H, which can normally obtain power from the non-failed first battery system 111, and the second dc bus bar 121L may supply power via the first dc bus bar 121H or the other second dc bus bar 121L, which is electrically adjacent to the first dc bus bar 121H, according to the on-off of the third branch C3 between the first dc bus bars 121H.

In the case where all of the normal power supply branches of the plurality of first battery systems 111 fail, in this case, in an emergency situation, the power supply to the first dc bus bar 121H cannot be ensured, but in order to avoid complete runaway of the electric aircraft 10, the corresponding second dc bus bars 121L are supplied with power via the second battery systems 112, so that the operation of the fly-by-wire actuation units 141 connected to the respective second dc bus bars 121L is ensured. The power supply branch, i.e. the emergency power supply branch, is the second battery system 112

And a second dc bus bar 121L. In addition, in each switchboard box 121, the plurality of second dc bus bars 121L also have redundant power supply capability, in order to ensure that the corresponding second dc bus bars 121H can work normally when there is a partial failure of the emergency power supply branch of the plurality of sets of second battery systems 112, power is supplied to them by the adjacent non-failed second battery systems 112, that is, power is supplied to the second dc bus bars 121L corresponding thereto in the switchboard box 121 by the adjacent non-failed second battery systems 112 first; at the same time, current flows from the second DC bus 121L corresponding to the non-failed second battery system 112The corresponding second branch C2 supplies power to the second dc bus bar 121L (also referred to as "the second dc bus bar as the redundant power supply object") corresponding to the failed second battery system 112 to ensure the operation of the fly-by-wire actuation system 140 to the maximum. That is, when all of the normal power supply branches fail, the third power supply path or the fourth power supply path different from the first power supply path and the second power supply path is switched, and the second dc bus bar 121L may be supplied with power via the corresponding second battery system 112 or another second dc bus bar 121L electrically connected adjacent thereto.

In other words, each of the switchboard boxes 121 of the electric aircraft integrated electric propulsion system 100 of the present invention has a plurality of energizing channels using the first battery system 111 as a power supply source inside, and in the case of normal power supply, each of the plurality of first direct current bus bars 121H is respectively supplied with power by the respective directly connected first battery system 111, and each of the plurality of second direct current bus bars 121L also respectively obtains power from the electrically adjacent connected first direct current bus bars 121H via the directly connected voltage conversion device 121a, and when a fault occurs in a certain energizing channel, the first direct current bus bar 121H corresponding to the first battery system 111 in the energizing channel is switched to be supplied with power by the adjacent energizing channel. When an emergency situation occurs, the plurality of second battery systems 112 provide emergency power to the fly-by-wire actuation unit 141 on the plurality of second dc bus bars 121L, and each of the plurality of second dc bus bars 121L's ability to specifically supply power redundantly can be powered by the plurality of second battery systems 112.

Stated another way, each first dc bus bar 121H of the plurality of first dc bus bars 121H (also referred to as a "target first dc bus bar") allows for a power supply path to be selected by the first battery system 111 directly connected to the target first dc bus bar and the first dc bus bar electrically connected adjacent to the first dc bus bar 121H (i.e., the target first dc bus bar). Each of the plurality of second dc bus bars 121L (also referred to as "target second dc bus bars") allows one power supply path to be selected by the second battery system 112 directly connected to the target second dc bus bar, the first dc bus bar electrically connected adjacent to the second dc bus bar 121 (i.e., target second dc bus bar), and the second dc bus bar electrically connected adjacent to the second dc bus bar 121 (i.e., target second dc bus bar).

In the case where the first dc bus bar and/or the second dc bus bar electrically adjacent to the target first dc bus bar or the target second dc bus bar cannot be directly connected (supplied with power) from the first battery system 111 or the second battery system 112, the first dc bus bar and/or the second dc bus bar to be connected next electrically adjacent thereto is selected again.

Additional advantages and modifications will readily appear to those skilled in the art. Therefore, the utility model in its broader aspects is not limited to the specific details and representative embodiments shown and described herein. Accordingly, modifications may be made without departing from the spirit or scope of the general inventive concept as defined by the appended claims and their equivalents.

In the electric aircraft integrated electric propulsion system 100 of the present utility model, the number of the switchboard boxes 121, the first battery system 111, the second battery system 112, the electric propulsion units 131, the fly-by-wire actuation units 141, the first dc bus bars 121H, the second dc bus bars 121L, and the like is not limited to the number shown in fig. 2.

For example, in the electric aircraft integrated electric propulsion system 100 according to the present utility model, only one switchboard box 121 may be provided on each of the left and right wings 11, and two or more switchboard boxes 121 may be provided. The number of the distribution boxes 121 provided on the left and right wings 11 is preferably the same, but may be different, for example, one distribution box 121 may be provided on one wing 11 and two distribution boxes 121 may be provided on the other wing 11.

In addition, for one switchboard 121, there may be two or more first battery systems 111 and two or more second battery systems 112, and each first battery system 111 may be provided with two or more sets of electric propulsion units 131, and each second battery system 112 may be provided with two or more sets of electric flight control units 141.

In the case where a plurality of switchboard boxes 121 are arranged on the wing 11 on one side, it is not necessarily required that each switchboard box 121 has the above-described structure as long as at least one switchboard box 121 has the above-described structure.

In addition, in the structure of the switchboard 121 shown in fig. 4, there are two first direct current bus bars 121H and two second direct current bus bars 121L, which are configured to: one first dc bus bar 121H (e.g., upper left in fig. 4) is connected to another first dc bus bar 121H (e.g., upper right in fig. 4) that is electrically adjacent via a first leg C1; a first dc bus bar 121H (e.g., upper left in fig. 4) is connected to a second dc bus bar 121L (e.g., lower left in fig. 4) that is electrically adjacent thereto via a third leg C3; one second dc bus bar 121L (e.g., lower left in fig. 4) is connected to another second dc bus bar 121L (e.g., lower right in fig. 4) that is electrically adjacent through a second leg C2; the other first dc bus bar 121H (for example, upper right in fig. 4) is connected to the electrically adjacent second dc bus bar 121L (for example, lower right in fig. 4) through the third branch C3, but in the structure of the switchboard box 121 of the electric aircraft integrated electric propulsion system 100 of the present invention, three or more first dc bus bars 121H and three or more second dc bus bars 121L may be provided, and as an example, three or more cases may be configured such that they are configured such that: one first dc bus bar 121H (e.g., upper left) is connected to an electrically adjacent (e.g., upper middle) further first dc bus bar 121H by a first leg C1, and the further first dc bus bar 121H is also connected to an electrically adjacent (e.g., upper right) further first dc bus bar 121H by a first leg C1; one (e.g., lower left) second dc bus bar 121L is connected to an electrically adjacent (e.g., lower middle) second dc bus bar 121L by a second leg C2, and the other second dc bus bar 121L is also connected to an electrically adjacent (e.g., lower right) further second dc bus bar 121L by a second leg C2; one (e.g., upper left) first dc bus bar 121H is connected to an electrically adjacent (e.g., lower left) second dc bus bar 121L by a third leg C3, the other (e.g., upper middle) first dc bus bar 121H is connected to an electrically adjacent other (e.g., lower middle) second dc bus bar 121L by a third leg C3, and the other (e.g., upper right) first dc bus bar 121H is connected to an electrically adjacent other (e.g., lower right) second dc bus bar 121L by a third leg C3.

Here, in the specification of the present application, the expressions "a specific component a electrically adjacent to a specific component B", "a specific component a is connected to a specific component B electrically adjacent to (and a specific component a)" and the like mean that no other component a or component B is present in the branch between the specific component a and the specific component B, for example, in fig. 4, the upper left first dc bus bar 121H is electrically adjacent to the lower left second dc bus bar 121L and to the upper right first dc bus bar 121H, but is not electrically adjacent to the lower right second dc bus bar 121L. In addition, a component C other than the components a and B may be provided in a branch between the specific component a and the specific component B, and for example, in fig. 4, a voltage conversion device 121a (and a circuit breaker) or the like may be provided between the first dc bus 121H on the upper left and the second dc bus 121L on the lower left, which are electrically connected adjacently. In this case, the component a and the component B which are electrically connected to each other are not necessarily two components which are physically adjacent to each other or are close to each other, and may be two components which are physically far apart, so long as the above situation is satisfied.

In addition, in the system configuration diagram of the electric aircraft integrated electric propulsion system shown in fig. 3, in order to avoid complication of the wiring, the switchboard boxes 121 on the wing 11 located on the left side (for example, the upper side in fig. 3) and the switchboard boxes 121 on the wing 11 located on the right side (for example, the lower side in fig. 3) are electrically connected to each other in such a manner that only one of the respective first dc bus bars 121H is connected through the electric wiring CA and only one of the respective second dc bus bars 121L is connected through the electric wiring CB, but the present invention is not limited thereto, and may be configured such that each of the first dc bus bars 121H is connected correspondingly through the electric wiring CA and each of the second dc bus bars 121L is connected correspondingly through the other electric wiring CB.

In the present invention, 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, but the present invention is not limited to this, and the first battery system 111 and the second battery system 112 may be any battery system suitable for use in an electric aircraft. Further, considering the first battery system 111 as a main power source and the second battery system 112 as an emergency (standby) power source, it is preferable that the capacity of the first battery system 111 is larger than that of the second battery system 112. Meanwhile, in the embodiment of the specification, the second battery system 112 is exemplified as a chargeable/dischargeable battery system, but the second battery system 112 is not necessarily provided with a charging function, and may be a discharge-only battery system.

Claims (13)

1. An electric aircraft comprehensive electric propulsion system is characterized in that,

the electric aircraft integrated electric propulsion system is independent of an on-board electric energy system which is connected with a plurality of electrified loads in a hanging way,

the electric aircraft comprehensive electric propulsion system combines a distributed electric propulsion system integrated on the wing of the electric aircraft and an fly-by-wire actuation system into a whole and is connected with a unified power supply,

in normal power supply, a first power supply circuit is used for supplying power to the distributed electric propulsion system and the fly-by-wire actuation system, and

and when the normal power supply is in a fault condition of partial failure or in an emergency condition of total failure, switching a path for supplying power to the distributed electric propulsion system and the fly-by-wire control actuation system into a power supply path different from the first power supply path through an energy management part.

2. An electric aircraft integrated electric propulsion system according to claim 1,

the energy management section is a plurality of switchboard boxes,

at least one of the plurality of switchboard boxes comprises a plurality of first direct current bus bars connected in series and a plurality of second direct current bus bars connected in series, wherein:

among the plurality of first direct current bus bars connected in series, every two of the first direct current bus bars which are electrically adjacent are connected with each other through a first branch;

Among the plurality of second direct current bus bars connected in series, every two of the second direct current bus bars which are electrically adjacent are connected with each other through a second branch;

the plurality of first dc bus bars and the plurality of second dc bus bars are connected to each other by a third branch provided with a voltage converting device,

the plurality of distribution boards are connected to each other by an electrical line corresponding to at least one of the first dc bus bars, and connected to each other by another electrical line corresponding to at least one of the second dc bus bars, thereby forming the energy management unit.

3. An electric aircraft integrated electric propulsion system according to claim 2,

the unified power supply includes a plurality of first battery systems and a plurality of second battery systems,

each set of the first battery systems is correspondingly connected to each first direct current bus bar,

each set of the second battery systems is correspondingly connected to each second direct current bus bar.

4. An electric aircraft integrated electric propulsion system according to claim 3,

the first power supply path for supplying power to the distributed electric propulsion system and the fly-by-wire actuation system is as follows:

Power is supplied to the corresponding first dc bus bars in the switchboard box by each set of the first battery system to ensure operation of the electric propulsion units constituting the distributed electric propulsion system connected to each first dc bus bar, and then current flows into each second dc bus bar connected to each first dc bus bar in an electrically adjacent manner via the corresponding third branch to ensure operation of the electric flight control actuation units constituting the electric flight control actuation system connected to each second dc bus bar.

5. An electric aircraft integrated electric propulsion system according to claim 3,

when the normal power supply to each first direct current bus bar is partially failed, namely, when a plurality of sets of normal power supply branches of the first battery system have partial failure, a second power supply path for supplying power to the distributed electric propulsion system and the fly-by-wire actuation system is as follows:

supplying power to a corresponding first dc bus bar in the switchboard box through an adjacent first battery system which is not failed to ensure the operation of the electric propulsion units which are connected with the first dc bus bar and form the distributed electric propulsion system, then current flows into a second dc bus bar which is connected with the first dc bus bar in an electric adjacent way through a corresponding third branch to ensure the operation of a fly-by-wire actuation unit which is connected with the second dc bus bar and forms the fly-by-wire actuation system,

And simultaneously, current is supplied from a first direct current bus bar corresponding to a first battery system which does not fail to a first direct current bus bar corresponding to the first battery system which fails, namely, a first direct current bus bar which is a redundant power supply object, so as to ensure the operation of an electric propulsion unit connected with the first direct current bus bar which is the redundant power supply object, and then, current flows from the first direct current bus bar which is the redundant power supply object to a second direct current bus bar which is electrically adjacent to the first direct current bus bar through a corresponding third branch, or flows from the second direct current bus bar which is electrically adjacent to the first direct current bus bar which is the redundant power supply object through a corresponding second branch.

6. An electric aircraft integrated electric propulsion system according to claim 3,

when all normal power supply to each first direct current bus bar fails, namely, in an emergency situation that all normal power supply branches of a plurality of sets of first battery systems fail, a third power supply path for supplying power to the fly-by-wire actuation system is as follows:

and supplying power to the corresponding second direct current bus bars through the second battery system so as to ensure the operation of fly-by-wire actuation units connected with the second direct current bus bars and forming the fly-by-wire actuation system.

7. An electric aircraft integrated electric propulsion system according to claim 6,

when a plurality of emergency power supply branches of the second battery system have partial failure, a fourth power supply path for supplying power to the fly-by-wire control actuation system is as follows:

power is supplied to its corresponding second dc bus bar in the switchboard box by an adjacent non-failed second battery system,

and simultaneously, current is supplied from the second direct current bus bar corresponding to the second battery system which does not fail to the second direct current bus bar corresponding to the second battery system which fails, namely the second direct current bus bar which is the redundant power supply object through the corresponding second branch circuit.

8. An electric aircraft integrated electric propulsion system according to claim 3,

each of the plurality of first dc bus bars allows one power supply path to be selected by the first battery system and the first dc bus bar that is electrically adjacent to the first dc bus bar.

9. An electric aircraft integrated electric propulsion system according to claim 8,

each of the plurality of second dc bus bars allows a power supply path to be selected by the second battery system, a first dc bus bar electrically connected adjacent to the second dc bus bar, and a second dc bus bar electrically connected adjacent to the second dc bus bar.

10. An electric aircraft integrated electric propulsion system according to claim 3,

the electric aircraft comprehensive electric propulsion system is characterized in that a plurality of power-on channels taking the first battery system as a main power supply are arranged in each distribution board box,

in the case of normal power supply, each of the plurality of first direct current bus bars is supplied with power by the respective directly connected first battery system, each of the plurality of second direct current bus bars also obtains power from the electrically adjacent connected first direct current bus bars via the directly connected voltage conversion device,

when a fault condition occurs in a certain power supply channel, the first direct current bus bar corresponding to the first battery system in the power supply channel is switched to be supplied by an adjacent power supply channel,

when an emergency occurs, the plurality of second battery systems serve as standby power sources for providing emergency power sources for fly-by-wire actuation units on the plurality of second direct current bus bars, and the capability of each of the plurality of second direct current bus bars to be specifically and redundantly powered can be powered by the plurality of second battery systems.

11. An electric aircraft integrated electric propulsion system according to any one of claims 3 to 10,

The electric aircraft comprehensive electric propulsion system is symmetrically arranged on the wings on the left side and the right side of the electric aircraft in a left-right symmetry manner,

each wing on each side comprises a distribution board box, four sets of electric propulsion units forming the distributed electric propulsion system, four sets of fly-by-wire control actuation units forming the fly-by-wire control actuation system, and two sets of first battery systems and two sets of second battery systems serving as a unified power supply source.

12. An electric aircraft integrated electric propulsion system according to claim 11,

the first battery system is a high specific energy lithium battery system,

the second battery system is an emergency lithium battery system.

13. An electric aircraft integrated electric propulsion system according to claim 12,

the emergency lithium battery system is provided with a DC/DC converter capable of charging and discharging.

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