CN115432190A - Electric aircraft comprehensive electric propulsion system - Google Patents
Electric aircraft comprehensive electric propulsion system Download PDFInfo
- Publication number
- CN115432190A CN115432190A CN202211158972.4A CN202211158972A CN115432190A CN 115432190 A CN115432190 A CN 115432190A CN 202211158972 A CN202211158972 A CN 202211158972A CN 115432190 A CN115432190 A CN 115432190A
- Authority
- CN
- China
- Prior art keywords
- electric
- power supply
- bus bar
- electric propulsion
- propulsion system
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
Images
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
Abstract
The comprehensive electric propulsion system of the electric airplane can integrate a distributed electric propulsion system and an electric transmission control actuation system which are integrated on a wing, provide reliable, efficient and stable flight power for the electric airplane, and simultaneously reduce the weight and energy loss of the system. The electric airplane comprehensive electric propulsion system is independent of an airborne electric energy system for hanging a plurality of electrified loads, integrates a distributed electric propulsion system and an electric transmission flight control actuation system integrated on wings of an electric airplane into a unified power supply source and is connected with the unified power supply source, when the electric airplane comprehensive electric propulsion system supplies power to the distributed electric propulsion system and the electric transmission flight control actuation system through a first power supply path during normal power supply, and when the normal power supply has a fault condition of partial failure or an emergency condition of complete failure, a path for supplying power to the distributed electric propulsion system and the electric transmission flight control actuation system is switched into a power supply path different from the first power supply path through an energy management part.
Description
Technical Field
The invention relates to the electric propulsion technology of an airplane, in particular to an electric airplane comprehensive electric propulsion system.
Background
With the development of the electrification technology, the electric aircraft adopts a battery and a propulsion motor to replace the driving mode of fuel oil and an engine of the traditional aircraft, so that the noise and the emission are reduced, and meanwhile, the energy utilization rate is improved. At present, the limitation of motor technology and power distribution technology is adopted, and a common practice is to arrange a plurality of electric propulsion units at the edge of the wings of the airplane and integrate the electric propulsion units into a distributed electric propulsion system, so that effective thrust is provided for the electric airplane under the condition of meeting the existing mature power electronic technology.
However, the moments generated by the thrust forces at various locations on the wing are different, and the use of only a distributed electric propulsion system can have a significant negative impact on the stability of the aircraft flight.
Meanwhile, the specific energy of the battery also limits the output capacity of the distributed electric propulsion system, and the influence of thrust output and flight attitude on the flight performance must be comprehensively considered, so that the limited electric power and energy resources are fully utilized, and the flight capacity of the airplane is improved.
In addition, the electric flight control system (also called electric flight control actuation system) has the characteristics of short-time and high-power charging and discharging, which causes the weight and the loss of the power transmission and distribution device of the upstream branch circuit to be far higher than those of the traditional flight control system.
Disclosure of Invention
Therefore, the present invention is made to solve the above problems of the prior art, and an object of the present invention is to provide an integrated electric propulsion system for an electric aircraft, which integrates a distributed electric propulsion system and an electric fly-by-wire actuation control system integrated on a wing, so as to provide reliable, efficient and stable flight power for the electric aircraft, and reduce the weight and energy loss of the system.
In order to achieve the above object, the present invention provides an electric aircraft integrated electric propulsion system independent of an onboard electric energy system for hooking a plurality of electrified loads, wherein the electric aircraft integrated electric propulsion system integrates a distributed electric propulsion system and an electric fly-by-wire control actuation system integrated on a wing of an electric aircraft into a unified power supply, and supplies power to the distributed electric propulsion system and the electric fly-by-wire control actuation system through a first power supply path during normal power supply, and switches a path for supplying power to the distributed electric propulsion system and the electric fly-by-wire control actuation system into a power supply path different from the first path through an energy management unit during a fault condition in which the normal power supply partially fails or an emergency condition in which all the normal power supply fails.
According to the above constitution, compared with the architecture of the prior art, the electric aircraft integrated electric propulsion system of the invention has the following beneficial effects:
(1) The comprehensive electric propulsion system of the electric airplane and the onboard electric energy system are mutually independent, the control is simple, the realization difficulty is low, and the realization and the modification are easy;
(2) The electric airplane comprehensive electric propulsion system has the characteristics of a distributed energy system, cables and conversion devices for equipment power supply are reduced, the system weight is reduced, the electric energy loss is reduced, and the reliability of the system is also improved;
(3) The comprehensive electric propulsion system of the electric airplane integrates two system functions of thrust output (a distributed electric propulsion system) and attitude control (an electric fly-by-wire control actuation system), provides a foundation for comprehensive and cooperative control of the thrust output (the distributed electric propulsion system) and the attitude control (the electric fly-by-wire control actuation system), and is favorable for improving the performance, the stability and the energy utilization rate of the electric airplane.
(4) The electric aircraft comprehensive electric propulsion system is connected with the distributed electric propulsion system and the fly-by-wire control actuation system by adopting the uniform power supply source, when a part of main power supply fails (fails), the non-failed main power supply can be utilized to supply power to the whole set of distributed electric propulsion system and the whole set of fly-by-wire control actuation system by using a power supply path different from a first power supply path during normal power supply, and two system functions of thrust output and attitude control of the electric aircraft are ensured to the maximum extent. Even if all the main power supplies fail (fail), even if a part of the standby power supplies fail (fail), the whole set of fly-by-wire control actuation system can still be supplied with power by the standby power supplies which do not fail, so that the safe landing of the electric airplane can be ensured through attitude control.
Preferably, the energy management unit is a plurality of switchboard boxes. The plurality of distribution board boxes may be symmetrically installed on the left and right wings, or the number of distribution board boxes installed on one wing may be different from the number of distribution board boxes installed on one wing.
In one example, for example, one distribution box is included on each side of the wing. Also, preferably, each of the distribution cabinets includes a plurality of first dc bus bars connected in series and a plurality of second dc bus bars connected in series. In the distribution board box of such a structure, of the plurality of first dc bus bars connected in series, two of the first dc bus bars electrically adjacent to each other are connected to each other through first branches, respectively; in the plurality of second direct current bus bars connected in series, every two electrically adjacent second direct current bus bars are connected with each other through a second branch circuit respectively; the plurality of first direct current bus bars and the plurality of second direct current bus bars are connected to each other through a third branch provided with a voltage conversion device, in correspondence with each other one by one. The energy management unit is configured such that one or more first dc bus bars of one of the distribution cabinets are connected to one or more first dc bus bars of the other distribution cabinet via an electrical line, and one or more second dc bus bars of one of the distribution cabinets are connected to one or more second dc bus bars of the other distribution cabinet via an electrical line.
Since a plurality of distribution board boxes can be designed with a plurality of 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 switchboard box on the wing on one side and more than two switchboard boxes on the wing on the other side. In the three or more distribution boxes on both sides, as long as at least one of the distribution boxes has the structure of the distribution box of the above-described one example, the three or more distribution boxes are connected to each other with the first dc bus bar through an electric line and the second dc bus bar through another electric line, thereby constituting the energy management section.
Further preferably, the first dc bus bar in each distribution box is connected to the first battery system, and the second dc bus bar in each distribution box is connected to the second battery system, and the first battery system and the second battery system constitute a unified power supply source.
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 connected electrically adjacent to the first dc bus bar. And, each of the plurality of second direct current bus bars allows one power supply path to be selected by the second battery system, the first direct current bus bar connected electrically adjacent to the second direct current bus bar, and the second direct current bus bar connected electrically adjacent to the second direct current bus bar.
According to the above configuration, each distribution board box of the electric aircraft integrated electric propulsion system has a plurality of power supply channels with the first battery system as a main power supply inside, and when an emergency situation occurs, the plurality of second battery systems serve as backup power supplies for the fly-by-wire control actuation units on the plurality of second dc bus bars.
In the case of a normal supply, each of the plurality of first dc busbars is supplied with power by the respective directly connected first battery system, and each of the plurality of second dc busbars also receives power from the electrically adjacently connected first dc busbars via the directly connected voltage conversion device. Therefore, when the power is normally supplied to each first dc bus, the first power supply path for supplying power to the distributed electric propulsion system and the electric fly-by-wire actuation system is: the power is supplied to the corresponding first direct current bus bars in the distribution board box through each set of first battery system so as to ensure the operation of the electric propulsion units which are connected with the first direct current bus bars and form the distributed electric propulsion system, and then the current flows into the second direct current bus bars which are electrically adjacent to the first direct current bus bars through the corresponding third branch circuits so as to ensure the operation of the electric fly-by-wire actuation units which are connected with the second direct current bus bars and form the electric fly-by-wire actuation system.
In addition, when a failure condition of a failure occurs in a certain power supply channel, the first dc bus bar corresponding to the first battery system in the energization channel is switched to be supplied with power from the adjacent energization channel. Therefore, in the event of a partial failure of the normal power supply to the respective first dc bus bars, i.e. in the event of a failure of partial failure of the normal power supply branches of the plurality of sets of first battery systems, the second power supply path for supplying power to the distributed electric propulsion system and the electric fly-by-wire actuation system is: the method comprises the steps that power is supplied to a first direct current bus bar corresponding to the first direct current bus bar in a distribution board box through an adjacent first battery system which does not fail so as to ensure the operation of an electric propulsion unit which is connected with the first direct current bus bar and forms a distributed electric propulsion system, then current flows into a second direct current bus bar which is electrically adjacent to the first direct current bus bar through a corresponding third branch so as to ensure the operation of an electric flight control actuating unit which is connected with the second direct current bus bar and forms an electric flight control actuating system, meanwhile, the current is supplied to the first direct current bus bar which corresponds to the first battery system which does not fail, namely the first direct current bus bar which is used as a redundant power supply object from the first direct current bus bar which corresponds to the first battery system which does not fail through a first branch so as to ensure the operation of the electric propulsion unit which is connected with the first direct current bus bar which is used as the redundant power supply object, then the current flows into the second direct current bus bar which is adjacent to the second direct current bus bar which is used as the redundant power supply object from the first direct current bus bar which is connected with the redundant power supply object through the second direct current bus bar which is connected with the second direct current bus bar or is connected with the second direct current bus bar which is adjacent to the redundant power supply object.
That is, as long as a certain first dc bus bar is not completely electrically disconnected from the first battery system and the first dc bus bar electrically connected adjacent to the first dc bus bar, the operation of the electric propulsion unit connected to the first dc bus bar can be ensured, whereby both the thrust output and attitude control system functions of the electric aircraft can be ensured to the maximum extent.
When all the normal power supplies to the first direct current bus bars fail, namely under the emergency condition that all the normal power supply branches of the multiple 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 electric fly-by-wire control actuating unit which is connected with each second direct current bus bar and forms the electric fly-by-wire control actuating system.
Preferably, each of the plurality of second dc bus bars is capable of being powered by a plurality of second battery systems. At this time, when there is partial failure in the emergency power supply branches of the multiple sets of second battery systems, the fourth power supply path for supplying power to the fly-by-wire actuation system is as follows: and the power is supplied to the corresponding second direct current bus bar in the distribution board box through the adjacent non-failed second battery system, and simultaneously, the current is supplied to the second direct current bus bar which is a redundant power supply object and corresponds to the failed second battery system from the second direct current bus bar which corresponds to the non-failed second battery system through the corresponding second branch, so that the operation of the electric flight control actuating system is ensured to the maximum 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 adjacent to the second dc bus bar, and the second dc bus bar electrically connected adjacent to the second dc bus bar, the operation of the electrically fly actuated unit connected to the second dc bus bar can be ensured. Therefore, the two system functions of attitude control of the electric airplane can be guaranteed to the maximum extent, and the risk of the electric airplane out of control under the emergency condition is reduced.
In one embodiment of the invention, the integrated electric propulsion system of the electric airplane is symmetrically arranged on the left and right wings of the electric airplane, 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 fly control actuation units forming an electric fly control actuation system, and two sets of first battery systems and two sets of second battery systems serving as unified power supplies. However, it should be understood that this is only for a particular architecture of a certain type or 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 and discharge function, or may be a battery system having only a discharge function. In the former case, it is preferable to have a chargeable/dischargeable DC/DC converter.
Drawings
FIG. 1 is a block schematic diagram of an electric aircraft integrated electric propulsion system of the present invention.
Fig. 2 is a schematic physical configuration of an electric aircraft integrated electric propulsion system of the present invention disposed on an electric aircraft.
Fig. 3 is a system architecture diagram of an electric aircraft integrated electric propulsion system of the present invention disposed on an electric aircraft.
Fig. 4 is a system architecture diagram of a wing of the electric aircraft integrated electric propulsion system of the present invention on one side, in which the specific configuration of the inside of the switchboard cabinet is shown.
(symbol description)
10. An electric airplane;
11. an airfoil;
100. an electric aircraft integrated electric propulsion system;
110. a power supply source;
111. a first battery system;
111a lithium battery pack;
111b power conversion devices;
112. a second battery system;
112a lithium battery pack;
112b DC/DC converter;
120. an energy management unit;
121. a distribution board cabinet;
121a voltage conversion device;
121H first dc bus bar;
a 121L 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 a propulsion device;
140. an fly-by-wire actuation system;
141. an fly-by-wire actuation unit;
141a three-phase inverter;
141b a servo motor;
141c actuating cylinder;
c1 A first branch;
c2 A second branch;
c3 A third branch;
CA. A CB electrical line.
Detailed Description
Hereinafter, specific 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 of the present invention is independent of an electrified onboard system, i.e., an onboard electric energy system, to which a plurality of electrified loads are hooked, and as shown in fig. 1, a distributed electric propulsion system 130 and an electric flight control actuation system 140 integrated on a wing 11 of an electric aircraft 10 are integrated to be connected to a unified power supply 110, and a unified and mutually coordinated power energy supply is provided for the flight of the electric aircraft through an energy management unit 120, and at the same time, the weight and energy consumption of the electric aircraft can be effectively reduced.
The energy management section 120 manages, distributes, and transmits electric energy of the unified power supply source 110 by a plurality of distribution board boxes 121 described later, thereby supplying power to the distributed electric propulsion system 130 and the electric flight control actuation system 140. The distributed electric propulsion system 130 uses a plurality of sets of electric propulsion units 131, described later, as a thrust generation device to provide thrust for aircraft flight. The fly-by-wire actuation system 140 is used to change the tilt angles of ailerons, flaps, etc. by using a plurality of sets of fly-by-wire actuation units 141, which will be described later, to provide different flight attitudes for the aircraft.
Fig. 2 shows a physical structure of the electric-powered aircraft integrated electric propulsion system 100 of the invention disposed on the electric-powered aircraft 10, and fig. 3 is a system architecture diagram of the electric-powered aircraft integrated electric propulsion system 100 of the invention disposed on the electric-powered 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 sides of the wing 11 of the electric aircraft 10, and includes at least one (one shown in fig. 2) distribution board box 121, a plurality of sets (four sets shown in fig. 2) of electric propulsion units 131 constituting the distributed electric propulsion system 130, and a plurality of sets (four sets shown in fig. 2) of fly-by-wire control actuation units 141 constituting the fly-by-wire control actuation system 140 on each side of the wing 11, respectively, as shown in fig. 3, the plurality of sets of electric propulsion units 131 and the plurality of sets of fly-by-wire control actuation units 141 on each side of the wing 11 are connected to the distribution board box 121 on that side.
The sets of electric propulsion units 131 are symmetrically arranged left and right on the front side edge of the wing 11 and are supplied with power by first (e.g., 270V high voltage) dc bus bars 121H in the respective distribution cabinets 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 uses a propeller or a ducted fan to convert the rotational speed and torque output by the motor into thrust.
In addition, the multiple sets of fly-by-wire actuation units 141 are symmetrically disposed left and right on the rear side edge of the wing 11, and are powered by the second (e.g., 28V low-voltage) dc bus bar 121L in the respective distribution board 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 constant rotational speed in accordance with a control command from the three-phase inverter 141 a. The actuator cylinder 141c is controlled to deflect a control surface of an aileron, a flap, or the like, thereby adjusting a flight attitude.
It should be noted that in the illustrated embodiment of the present invention, the electric airplane 10 has a smaller overall structure, so that the electric fly control actuation unit 141 is powered by, for example, a 28V low-voltage dc bus, and when the electric airplane 10 is larger, the voltage level of the power bus of the electric fly control actuation unit 141 will rise, for example, by a 270V high-voltage dc bus, depending on the specific size of the electric airplane 10 and the power of the electric fly control actuation unit 141.
In addition, a plurality of sets (two sets are shown in fig. 2) of the first battery systems 111 and a plurality of sets (two sets are shown in fig. 2) of the second battery systems 112 are respectively included on the wing 11 on each side, and as shown in fig. 3, the plurality of sets of the first battery systems 111 and the plurality of sets of the second battery systems 112 on the wing 11 on each side are respectively connected to the distribution board box 121. The first battery systems 111 and the second battery systems 112 on the left and right wing 11 are mounted in a left-right symmetrical manner to constitute the unified power supply 110 shown in fig. 1 for supplying power to the left and right distribution board boxes 121.
As shown in fig. 3, the first battery system 111 is, for example, a high specific energy lithium battery system having a lithium battery pack 111a and a power converting device 111b.
Although not shown, the lithium battery pack 111a includes a plurality of lithium battery cells and two independent sets of battery management systems, thermal management systems, housings, relays, sensors, and the like. Each lithium battery unit is composed of one or more lithium battery units. The two sets of battery management systems monitor the state of the lithium battery unit 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 working temperature of the lithium battery pack 111a is within a proper range. The lithium battery shell has the functions of fire prevention, explosion prevention and the like. The relay can open and cut off the high-voltage electric loop of the lithium battery unit, and external on-off is realized.
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 demand for flight power.
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 chargeable and dischargeable DC/DC converter 112b for supplying power to critical devices of the integrated electric propulsion system in an emergency situation. When the first battery system 111 is used as the power supply source of the electric aircraft integrated electric propulsion system 100, the lithium battery pack 112a can be charged by switching the DC/DC converter 112b to the charging mode, and the trickle charge can be performed when the second battery system 112 is close to full charge. On the other hand, when the second battery system 112 is used as the power supply source of the electric-powered aircraft integrated electric propulsion system 100 to supply power, the DC/DC converter 112b can be switched to the discharge mode to output the electric power of the lithium battery pack 112 a.
In addition, as shown in fig. 3, the electric aircraft integrated electric propulsion system 100 is configured such that at least one first dc bus bar 121H and at least one second dc bus bar 121L of each of the distribution board boxes 121 located on the left and right wings 11 are connected by the electric lines CA and CB, and the distribution board boxes 121 located on the left and right wings 11 form the energy management section 120 for managing, distributing, and transmitting the electric energy of the power supply source 110 (the plurality of sets of first battery systems 111 and the plurality of sets of second battery systems 112) shown in fig. 1, so that the power supply of each corresponding set of electric propulsion units 131 and each set of electric flight control actuation units 141 can be supplied by the control of the distribution board box 121 on the corresponding side by the unified power supply source 110, thereby providing a power supply path for multiple redundant power supplies and improving the fault tolerance of the power supply source 110.
In addition, the electric propulsion units 131 on the wings 11 on both left and right sides constitute the distributed electric propulsion system 130 shown in fig. 1. The fly-by-wire operation control units 141 on the left and right wings 11 constitute the fly-by-wire operation control system 140 shown in fig. 1.
Fig. 4 is a system architecture diagram of the wing 11 on one side of the electric aircraft integrated electric propulsion system 10 of the present invention, in which the specific configuration inside the distribution board box 121 is shown.
More specifically, as shown in fig. 4, each distribution board 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, two electrically adjacent ones of the plurality of first dc bus bars 121H are connected to each other by first branches C1, respectively, two electrically adjacent ones of the plurality of second dc bus bars 121L connected in series are connected to each other by second branches C2, respectively, between two electrically adjacent ones of the plurality of second dc bus bars 121L are connected to each other by third branches C3 provided with voltage conversion devices 121a, respectively, and in addition, the plurality of first dc bus bars 121H and the plurality of second dc bus bars 121L are connected to each other in correspondence with each other one by one. The voltage conversion device 121a is, for example, a DC/DC converter, and converts the direct current of the voltage supplied from the first direct current bus bar 121H into the direct current of the voltage required for the second direct current bus bar 121L.
In addition, a plurality of sets of the first battery systems 111 are respectively connected to the plurality of first dc bus bars 121H in each distribution board box 121, and each of the first dc bus bars 121H in each distribution board box 121 is connected to a plurality of sets (two sets shown in fig. 4) of the electric propulsion units 131. Further, a plurality of sets of the second battery systems 112 are respectively connected to the plurality of second dc bus bars 121L in each distribution board box 121, and each second dc bus bar 121L in each distribution board box 121 is connected to a plurality of sets (two sets are shown in fig. 4) of the electrically fly actuated units 141.
Further, although not shown, each distribution board box 121 further includes a switching device such as a bus bar contactor, a control relay, and a thermal breaker. 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 breaker. The distribution board box 121 is internally provided with a redundant fault-tolerant power grid structure for ensuring safe and reliable operation of the power grid of the electric aircraft integrated electric propulsion system 100.
In each distribution board box 121, when power is normally supplied, power is supplied to the corresponding first dc bus bar 121H in the distribution board box 121 through each set of the first battery system 111 to ensure thatThe electric propulsion unit 131 connected to each first dc bus bar 121H operates, and then the electric current flows into the second dc bus bar 121L electrically adjacent to each first dc bus bar via the corresponding third branch C3, so as to ensure the operation of the electric flight control actuation unit 141 connected to each second dc bus bar 121L. The power supply path at this time, i.e., the first power supply path, is
That is, in the case of normal power supply, each first direct current bus bar 121H is directly supplied with power from each set of the first battery system 111, and each second direct current bus bar 121L is supplied with power from the respective first direct current bus bars 121H electrically adjacent thereto via the voltage conversion device 121 a.
In addition, in each distribution board box 121, the plurality of first dc bus bars 121H have a redundant power supply capability, and when there is a partial failure in the normal power supply branches of the plurality of sets 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 normally operate, power is supplied thereto 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 distribution board box 121 to ensure 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 connected adjacent to the first dc bus bar 121H through the corresponding third branch C3 to ensure operation of the electric flight control 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 (also referred to as "first dc bus bar to be redundantly supplied") to the first dc bus bar 121H corresponding to the first battery system 111 that has failed via the corresponding first branch C1, so as to ensure operation of the electric propulsion unit 131 connected to the first dc bus bar 121H (i.e., the first dc bus bar 121H to be redundantly supplied), and then, if the third branch C3 between the first dc bus bar 121H to be redundantly supplied and the second dc bus bar 121L (also referred to as "second dc bus bar to be supplied") that is electrically adjacent to the first dc bus bar 121H to be redundantly supplied is a through-path (circuit breaker closed), current flows from the first dc bus bar to be redundantly supplied to the second dc bus bar 121L to be supplied via the corresponding third branch C3, and conversely, if the second dc bus bar 121H to be redundantly supplied and the second dc bus bar 121L to be redundantly supplied are electrically adjacent to the through-path (circuit breaker closed), current cannot flow from the second dc bus bar 121L to be redundantly supplied via the corresponding second dc bus bar 121L to be redundantly supplied (circuit breaker closed) and the second dc bus bar 121L to be electrically adjacent to be redundantly supplied via the second dc bus bar 121L to be redundantly supplied, and then, the second dc bus bar 121L to be connected to the second dc circuit breaker 121L to be connected to the second dc bus bar 121H via the corresponding second dc bus bar 121L. That is, in the case where there is a partial failure in the normal power supply branch, the first dc bus bar 121H that cannot normally obtain power from the failed first battery system 111 is switched to the second power supply path different from the first power supply path, and power can be indirectly obtained from the other first dc bus bar 121H that can normally obtain power from the non-failed first battery system 111, while the second dc bus bar 121L can supply power via the first dc bus bar 121H or the other second dc bus bar 121L that is electrically adjacent thereto, according to the on/off of the third branch C3 between the first dc bus bars 121H that are electrically adjacent thereto.
In the event of a total failure of the normal power supply branches of the multiple sets of first battery systems 111, which is an emergency situation at this time, the power supply to the first dc bus bars 121H cannot be guaranteed, but in order to avoid a complete runaway of the electric aircraft 10, the power is supplied to the corresponding second dc bus bars 121L by the second battery systems 112, so as to ensure the operation of the fly-by-wire actuation unit 141 connected to each second dc bus bar 121L. The power supply branch, namely the emergency power supply branch at the moment is
In addition, in each distribution board box 121, the plurality of second dc bus bars 121L also have the capability of redundant power supply, and when there is a partial failure in the emergency power supply branches of the plurality of sets of second battery systems 112, in order to ensure that the corresponding second dc bus bars 121H can work normally, power is supplied to the corresponding second dc bus bars 121L through the adjacent non-failed second battery systems 112, that is, power is first supplied to the corresponding second dc bus bars 121L in the distribution board box 121 through the adjacent non-failed second battery systems 112; at the same time, the current is supplied from the second dc bus bar 121L corresponding to the non-failed second battery system 112 to the second dc bus bar 121L corresponding to the failed second battery system 112 (also referred to as a "second dc bus bar to be redundantly supplied") via the corresponding second branch C2, so as to ensure the operation of the electric flight control actuation system 140 to the maximum extent. That is, when all of the normal power supply branches fail, the power supply branches are switched to a third power supply path or a fourth power supply path different from the first power supply path and the second power supply path, and the second dc bus bar 121L can 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 distribution board box 121 of the electric aircraft integrated electric propulsion system 10 of the present invention has a plurality of power supply channels inside, each of which is powered by the first battery system 111, and in the case of normal power supply, each of the plurality of first dc bus bars 121H is respectively powered by the respective directly connected first battery system 111, and each of the plurality of second dc bus bars 121L is also respectively powered from the electrically adjacent connected first dc bus bars 121H via the directly connected voltage conversion device 121a, and when a fault occurs in a certain power supply channel, the first dc bus bar 121H corresponding to the first battery system 111 in the power supply channel is switched to be powered by the adjacent power supply channel. When an emergency situation occurs, the plurality of second battery systems 112 provide emergency power to the electric fly-control actuation unit 141 on the plurality of second dc bus bars 121L, and the capability of each of the plurality of second dc bus bars 121L to supply power in a specific redundant manner can be provided by the plurality of second battery systems 112.
Put another way, each of the plurality of first dc bus bars 121H (also referred to as "target first dc bus bars") allows one 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 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).
When 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) from the first battery system 111 or the second battery system 112, the next electrically adjacent first dc bus bar and/or second dc bus bar is selected again.
Additional advantages and modifications will readily occur to those skilled in the art. Therefore, the invention 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-powered aircraft integrated electric propulsion system 100 of the present invention, the number of the distribution board boxes 121, the first battery system 111, the second battery system 112, the electric propulsion unit 131, the electric flight control actuation unit 141, the first dc bus bar 121H, the second dc bus bar 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 invention, only one distribution board box 121 may be provided on each of the left and right wings 11, and two or more distribution board boxes 121 may be provided. The number of distribution board boxes 121 disposed on the left and right wings 11 is preferably the same, but may be different, and for example, one distribution board box 121 may be disposed on one wing 11 and two distribution board boxes 121 may be disposed on the other wing 11.
In addition, one distribution board box 121 may have two or more first battery systems 111 and two or more second battery systems 112, and two or more electric propulsion units 131 may be provided for each first battery system 111, and two or more electric flight control operation units 141 may be provided for each second battery system 112.
In the case where a plurality of distribution board boxes 121 are disposed on one wing 11, it is not necessarily required that each distribution board box 121 has the above-described structure as long as at least one of the distribution board boxes 121 has the above-described structure.
In addition, in the configuration of the distribution board box 121 shown in fig. 4, there are two first dc bus bars 121H and two second dc bus bars 121L arranged such that: one (e.g., upper left in fig. 4) first dc bus bar 121H is connected to another electrically adjacent (e.g., upper right in fig. 4) first dc bus bar 121H through a first branch 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 branch C3; one second direct current bus bar 121L (e.g., lower left in fig. 4) is connected to another second direct current bus bar 121L that is electrically adjacent (e.g., lower right in fig. 4) by a second branch C2; the other (for example, the upper right one in fig. 4) first dc bus bar 121H and the electrically adjacent (for example, the lower right one in fig. 4) second dc bus bar 121L are connected by the third branch C3, but the distribution board box 121 of the electric aircraft integrated electric propulsion system 100 of the present invention may have three or more first dc bus bars 121H and three or more second dc bus bars 121L, and in the case of three, they may be arranged: one (e.g., upper left) first dc bus bar 121H is connected to another electrically adjacent (e.g., upper middle) first dc bus bar 121H via a first branch C1, and the other first dc bus bar 121H is also connected to another electrically adjacent (e.g., upper right) first dc bus bar 121H via a first branch C1; one (e.g., lower left) second dc bus bar 121L is connected to another electrically adjacent (e.g., lower middle) second dc bus bar 121L via a second branch C2, and the another second dc bus bar 121L is also connected to another electrically adjacent (e.g., lower right) second dc bus bar 121L via a second branch C2; one (e.g., upper left) first dc bus bar 121H is connected to one (e.g., lower left) electrically adjacent second dc bus bar 121L via a third branch C3, another (e.g., upper middle) first dc bus bar 121H is connected to another (e.g., lower middle) electrically adjacent second dc bus bar 121L via a third branch C3, and yet another (e.g., upper right) first dc bus bar 121H is connected to yet another (e.g., lower right) electrically adjacent second dc bus bar 121L via a third branch C3.
Here, in the specification of the present application, the expressions "a specific component a electrically connected to a specific component B", "a specific component a is connected to a specific component B electrically adjacent to (with the specific component a)" and the like mean that there is no other component a or component B in a branch between the specific component a and the specific component B, and for example, in fig. 4, the upper left first dc bus bar 121H, the lower left second dc bus bar 121L, and the upper right first dc bus bar 121H are electrically connected to each other in an adjacent manner, but are not electrically connected to the lower right second dc bus bar 121L. Further, there may be components C other than a component a and a component B in the branch between a specific component a and a specific component B, for example, in fig. 4, there may be a voltage conversion device 121a (and a circuit breaker) or the like between an upper left first dc bus bar 121H and a lower left second dc bus bar 121L that are electrically adjacently connected. In addition, the component a and the component B electrically connected to each other at this time do not necessarily mean two components physically adjacent to each other or closely adjacent to each other, and may be two components physically distant from each other as long as the above is satisfied.
In addition, in the system architecture diagram of the electric aircraft integrated electric propulsion system shown in fig. 3, in order to avoid complication of wiring, the distribution board box 121 on the wing 11 on the left side (for example, the upper side in fig. 3) and the distribution board box 121 on the wing 11 on the right side (for example, the lower side in fig. 3) are electrically connected to each other in such a manner that only each of the first dc bus bars 121H is connected by the electric wiring CA and only each of the second dc bus bars 121L is connected by the electric wiring CB, but the present invention is not limited thereto, and it may be configured such that each of the first dc bus bars 121H is correspondingly connected by the electric wiring CA and each of the second dc bus bars 121L is correspondingly connected by the other electric wiring CB.
In addition, 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 thereto, and the first battery system 111 and the second battery system 112 may be any battery system suitable for an electric aircraft. Further, considering that the first battery system 111 serves as a main power source and the second battery system 112 serves as an emergency (backup) power source, it is preferable that the first battery system 111 has a larger capacity than the second battery system 112. In the embodiment described above, the second battery system 112 is a chargeable and dischargeable battery system, but the second battery system 112 does not necessarily have to have a charging function and may be a battery system that is only discharged.
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 onboard electric energy system for hanging a plurality of electrified loads,
the electric airplane comprehensive electric propulsion system integrates a distributed electric propulsion system integrated on the wings of the electric airplane and an electric transmission flight control actuation system into a whole and is connected with a uniform power supply,
supplying power to the distributed electric propulsion system and the fly-by-wire actuation system with a first supply circuit during normal supply of power, and
and when the normal power supply has a fault condition of partial failure or an emergency condition of complete failure, switching a power supply path for supplying power to the distributed electric propulsion system and the electric fly-by-wire control actuation system into a power supply path different from the first power supply path through an energy management part.
2. The integrated electric propulsion system for electric aircraft of claim 1,
the energy management part is a plurality of switchboard boxes,
at least one of the distribution cubicles of the plurality of distribution cubicles comprises a plurality of first DC bus bars connected in series and a plurality of second DC bus bars connected in series, wherein:
in the plurality of first direct current bus bars connected in series, every two electrically adjacent first direct current bus bars are connected with each other through a first branch;
in the second direct current bus bars connected in series, every two electrically adjacent second direct current bus bars are connected with each other through a second branch circuit respectively;
the plurality of first direct current bus bars and the plurality of second direct current bus bars are connected to each other through a third branch provided with a voltage conversion device in correspondence one by one,
the plurality of distribution board boxes are configured such that at least one of the first dc bus bars is connected to each other by an electrical line, and at least one of the second dc bus bars is connected to each other by another electrical line, thereby configuring the energy management unit.
3. The integrated electric propulsion system for electric aircraft of claim 2,
the unified power supply comprises a plurality of sets of first battery systems and a plurality of sets of second battery systems,
each set of the first battery systems is connected to each first direct current bus bar correspondingly,
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 as claimed in claim 3,
when the power is normally supplied to each of the first dc bus bars, a first power supply path for supplying power to the distributed electric propulsion system and the fly-by-wire actuation system is:
supplying power to the corresponding first dc bus bars in the distribution box through each set of the first battery system to ensure operation of the electric propulsion units connected to the first dc bus bars that constitute the distributed electric propulsion system, and then supplying current to the second dc bus bars connected electrically adjacent to the first dc bus bars through the corresponding third branch paths to ensure operation of the electric flight control actuation units connected to the second dc bus bars that constitute the electric flight control actuation system.
5. An electric aircraft integrated electric propulsion system as claimed in claim 3,
in the event of a partial failure of the normal power supply to each of the first dc bus bars, i.e. in the event of a failure of partial failure of the normal power supply branches of the plurality of sets of the first battery system, the second power supply path for supplying power to the distributed electric propulsion system and the electric fly-by-wire actuation system is:
supplying power to a corresponding first dc bus bar in the distribution box via an adjacent, non-failed first battery system to ensure operation of an electric propulsion unit connected to the first dc bus bar, which constitutes the distributed electric propulsion system, and then flowing current via a corresponding third branch into a second dc bus bar connected electrically adjacent to the first dc bus bar to ensure operation of an electric flight control actuation unit connected to the second dc bus bar, which constitutes the electric flight control actuation system,
and simultaneously supplying current from a first dc bus corresponding to a first battery system that has not failed to the first battery system, that is, a first dc bus to be redundantly supplied with power, via the first branch to ensure operation of an electric propulsion unit connected to the first dc bus to be redundantly supplied with power, and then supplying current from the first dc bus to be redundantly supplied with power, via a corresponding third branch, to a second dc bus connected electrically adjacent to the first dc bus, or from the second dc bus to be electrically adjacent to the second dc bus via a corresponding second branch.
6. An electric aircraft integrated electric propulsion system according to claim 3,
when all the power supplies to the first direct current bus bars fail normally, namely in an emergency situation that all the normal power supply branches of the multiple sets of the 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 the fly-by-wire actuation unit which is connected with each second direct current bus bar and forms the fly-by-wire actuation system.
7. An electric aircraft integrated electric propulsion system as claimed in claim 6,
when partial failure exists in a plurality of sets of emergency power supply branches of the second battery system, a fourth power supply path for supplying power to the fly-by-wire actuation system is as follows:
the second dc bus bar in the distribution cabinet corresponding thereto is supplied with power through the adjacent non-failed second battery system,
and meanwhile, the current is supplied from the second direct current bus bar corresponding to the second battery system which is not failed to the second direct current bus bar corresponding to the second battery system which is failed, namely the second direct current bus bar which is a redundant power supply object, through the corresponding second branch.
8. An electric aircraft integrated electric propulsion system as claimed in 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 connected electrically adjacent to the first dc bus bar.
9. The integrated electric propulsion system for electric aircraft of claim 8,
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 connected electrically adjacent to the second dc bus bar, and the second dc bus bar connected electrically adjacent to the second dc bus bar.
10. An electric aircraft integrated electric propulsion system according to claim 3,
a plurality of electrifying channels which take the first battery system as a main power supply are arranged in each distribution board box of the electric aircraft integrated electric propulsion system,
in the case of a normal 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 is also supplied with power from the electrically adjacently connected first direct current bus bars via the directly connected voltage conversion device,
when a failure condition of a failure occurs in a certain power supply channel, the first dc bus bar corresponding to the first battery system in the energized channel is switched to be supplied with power from an adjacent energized channel,
when an emergency situation occurs, the second battery systems serve as backup power sources to provide emergency power for the electric fly-control actuation units on the second direct-current bus bars, and the power supply capacity of each of the second direct-current bus bars can be specifically and redundantly supplied by the second battery systems.
11. An electric aircraft integrated electric propulsion system according to any one of claims 3 to 10,
the electric airplane comprehensive electric propulsion system is symmetrically arranged on the left and right sides of the wings of the electric airplane,
the wings on each side respectively comprise a distribution board box, four sets of electric propulsion units forming the distributed electric propulsion system, four sets of electric transmission flight control actuating units forming the electric transmission flight control actuating system, and two sets of first battery systems and two sets of second battery systems serving as unified power supply sources.
12. The electric aircraft integrated electric propulsion system of 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.
Priority Applications (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211158972.4A CN115432190B (en) | 2022-09-22 | 2022-09-22 | Comprehensive electric propulsion system of electric aircraft |
| PCT/CN2023/085041 WO2024060567A1 (en) | 2022-09-22 | 2023-03-30 | Integrated electric propulsion system for electric aircrafts |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211158972.4A CN115432190B (en) | 2022-09-22 | 2022-09-22 | Comprehensive electric propulsion system of electric aircraft |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN115432190A true CN115432190A (en) | 2022-12-06 |
| CN115432190B CN115432190B (en) | 2023-06-02 |
Family
ID=84248757
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202211158972.4A Active CN115432190B (en) | 2022-09-22 | 2022-09-22 | Comprehensive electric propulsion system of electric aircraft |
Country Status (2)
| Country | Link |
|---|---|
| CN (1) | CN115432190B (en) |
| WO (1) | WO2024060567A1 (en) |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117439022A (en) * | 2023-12-20 | 2024-01-23 | 天津航空机电有限公司 | Distribution board box protection structure and protection method |
| WO2024060567A1 (en) * | 2022-09-22 | 2024-03-28 | 中国商用飞机有限责任公司 | Integrated electric propulsion system for electric aircrafts |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN204089325U (en) * | 2014-08-18 | 2015-01-07 | 江西洪都航空工业集团有限责任公司 | The novel power supply circuits of a kind of flight control system |
| CN108023399A (en) * | 2017-12-26 | 2018-05-11 | 北京航空航天大学 | A kind of neighbouring complementary power supply and distribution device of Helios |
| WO2019207256A1 (en) * | 2018-04-26 | 2019-10-31 | Safran | Hybrid propulsion system for an aircraft |
| EP3624301A1 (en) * | 2018-09-11 | 2020-03-18 | Embraer S.A. | Method and system for distributed electrical loads connected to shared power sources |
| US20220194614A1 (en) * | 2020-12-18 | 2022-06-23 | Airbus Operations Sas | Aircraft with electric or hybrid propulsion comprising a suitable electrical installation |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115432190B (en) * | 2022-09-22 | 2023-06-02 | 中国商用飞机有限责任公司 | Comprehensive electric propulsion system of electric aircraft |
-
2022
- 2022-09-22 CN CN202211158972.4A patent/CN115432190B/en active Active
-
2023
- 2023-03-30 WO PCT/CN2023/085041 patent/WO2024060567A1/en unknown
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN204089325U (en) * | 2014-08-18 | 2015-01-07 | 江西洪都航空工业集团有限责任公司 | The novel power supply circuits of a kind of flight control system |
| CN108023399A (en) * | 2017-12-26 | 2018-05-11 | 北京航空航天大学 | A kind of neighbouring complementary power supply and distribution device of Helios |
| WO2019207256A1 (en) * | 2018-04-26 | 2019-10-31 | Safran | Hybrid propulsion system for an aircraft |
| EP3624301A1 (en) * | 2018-09-11 | 2020-03-18 | Embraer S.A. | Method and system for distributed electrical loads connected to shared power sources |
| US20220194614A1 (en) * | 2020-12-18 | 2022-06-23 | Airbus Operations Sas | Aircraft with electric or hybrid propulsion comprising a suitable electrical installation |
Non-Patent Citations (1)
| Title |
|---|
| 戴泽华: "飞机供电系统集成设计方法研究" * |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO2024060567A1 (en) * | 2022-09-22 | 2024-03-28 | 中国商用飞机有限责任公司 | Integrated electric propulsion system for electric aircrafts |
| CN117439022A (en) * | 2023-12-20 | 2024-01-23 | 天津航空机电有限公司 | Distribution board box protection structure and protection method |
| CN117439022B (en) * | 2023-12-20 | 2024-04-09 | 天津航空机电有限公司 | Distribution board box protection structure and protection method |
Also Published As
| Publication number | Publication date |
|---|---|
| CN115432190B (en) | 2023-06-02 |
| WO2024060567A1 (en) | 2024-03-28 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| US11198515B2 (en) | Method and system for distributed electrical loads connected to shared power sources | |
| CN115432190B (en) | Comprehensive electric propulsion system of electric aircraft | |
| US7923857B2 (en) | System and method for supplying power for actuators on board an aircraft | |
| CN104884349B (en) | The management method of the electric power networks of aircraft | |
| CN107707011B (en) | Power supply and distribution device of solar aircraft | |
| CA2814564C (en) | Electrical power control system for a vehicle | |
| CN110797957A (en) | Dynamically configurable energy storage unit | |
| CN108023399B (en) | Solar aircraft adjacent complementary power supply and distribution device | |
| CN109660200B (en) | Solar unmanned aerial vehicle energy management system and solar unmanned aerial vehicle with same | |
| US11472297B2 (en) | Hybrid energy storage system | |
| CN112636459A (en) | Unmanned aerial vehicle intelligence power supply and distribution system | |
| CN113036894B (en) | Solar unmanned aerial vehicle electrical system architecture | |
| CN107681762A (en) | A kind of stratosphere aerostatics electric power-feeding structure | |
| CN211127291U (en) | Helicopter direct current distribution system | |
| US20230399099A1 (en) | Aircraft having a hybrid power source | |
| CN115765134A (en) | Aircraft unified power supply and distribution redundancy system | |
| CN115367122A (en) | Time-varying electrical load sharing in a power distribution network of an aircraft | |
| CN113794266A (en) | Distributed annular power distribution system architecture based on multi-bus configuration on satellite | |
| CN212085877U (en) | Large-scale unmanned aerial vehicle power supply redundancy system | |
| CN217374910U (en) | Hybrid power supply system and unmanned aerial vehicle | |
| CN113412219A (en) | Aircraft propulsion system | |
| US20220348348A1 (en) | Aircraft comprising an electric power distribution network | |
| RU2756390C1 (en) | Electric power supply system | |
| CN115663783A (en) | Bus power interactive power supply and distribution system based on minimum satellite system | |
| CN115912485A (en) | Power and airborne electric energy system of all-electric aircraft and related power supply system |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |