Power and on-board electrical energy systems for all-electric aircraft and related power supply systems

Abstract

The present disclosure relates to power and onboard electrical power systems for all-electric aircraft and related power supply systems. A power electrical power system includes a plurality of batteries for providing electrical power, a plurality of DC / DC converters, each of which is electrically connected to a respective one of the plurality of batteries to DC / DC convert direct current output by the battery, a plurality of distribution panels, each of which is electrically connected to an output of at least one of the plurality of DC / DC converters to receive DC / DC converted direct current, and a plurality of DC / AC inverters, each of which is electrically connected to an output of one of the plurality of distribution panels to convert direct current output from the distribution panel to three-phase alternating current, thereby providing electrical power to a drive device of the all-electric aircraft.

Description

Technical Field

[0001] This disclosure relates to the field of aviation electric propulsion, and more specifically to power and airborne electrical systems and related power systems for all-electric aircraft. Background Technology

[0002] Currently, in order to accelerate the achievement of goals such as energy conservation and emission reduction, reduce fuel consumption, and reduce aircraft weight, the all-electric design of aircraft has become an inevitable trend for future development.

[0003] Existing power system designs for all-electric aircraft still have shortcomings in terms of ease of implementation, control complexity, system reliability, weight, and efficiency.

[0004] This disclosure addresses, but is not limited to, the many factors mentioned above. Summary of the Invention

[0005] This disclosure proposes a discrete new energy aircraft power system. After comprehensively analyzing the advantages of the power system in terms of safety, weight, efficiency, and energy utilization, and considering the complexity of its implementation and the difficulty of its development, it provides an energy foundation for the design of all-electric aircraft. The aircraft power system disclosed herein proposes a fully electric aircraft power system with a discrete power system for the main power source and an independent power system for the onboard power source. Since the main power system and the onboard power system in this technical solution are two independent systems, the control scheme is simple, the implementation difficulty is low, and it is easy to retrofit existing aircraft by taking into account existing aircraft power systems and power systems. The discrete power conversion, transmission, and configuration links proposed in this technical solution are fewer, resulting in a higher overall system reliability.

[0006] According to a first aspect of this disclosure, a power system for an all-electric aircraft is provided, comprising: a plurality of batteries for providing power; a plurality of DC / DC converters, each of the plurality of DC / DC converters being electrically connected to a corresponding battery among the plurality of batteries to perform DC / DC conversion on direct current output from the battery; a plurality of distribution boxes, each of the plurality of distribution boxes being electrically connected to the output of at least one of the plurality of DC / DC converters to receive the DC / DC converted direct current; and a plurality of DC / AC inverters, each of the plurality of DC / AC inverters being electrically connected to the output of one of the plurality of distribution boxes to convert the DC output from the distribution box into three-phase alternating current, thereby providing power to the drive mechanism of the all-electric aircraft.

[0007] According to one embodiment, each of the plurality of distribution boxes includes at least a contactor, a control relay, a circuit breaker, and a busbar, wherein the busbar is electrically connected to the output of at least one of the plurality of DC / DC converters.

[0008] According to another embodiment, each of the plurality of distribution boxes also has a power controller for the status of the equipment inside the distribution box.

[0009] According to another embodiment, each of the plurality of distribution boxes is redundant with at least one of the other distribution boxes in the plurality of distribution boxes, such that if one of the plurality of distribution boxes fails, the distribution box that is redundant with the first one will take over the operation.

[0010] According to another embodiment, the plurality of distribution boxes includes a first distribution box and a second distribution box, the plurality of batteries includes a first battery and a second battery, the plurality of DC / DC converters includes a first DC / DC converter and a second DC / DC converter, the first battery is connected to the first distribution box via the first DC / DC converter, the second battery is connected to the second distribution box via the second DC / DC converter, wherein the first distribution box and the second distribution box are electrically connected together via a contactor.

[0011] According to another embodiment, each of the plurality of distribution boxes includes a first busbar and a second busbar, wherein the first busbar is electrically connected between one of the plurality of DC / DC converters and the second busbar, and the second busbar is electrically connected between the first busbar and one of the plurality of DC / AC inverters.

[0012] According to another embodiment, the plurality of distribution boxes includes a first, second, third, and fourth distribution box; the plurality of batteries includes a first, second, third, and fourth battery; the plurality of DC / DC converters includes a first, second, third, and fourth DC / DC converter; the first battery is connected to a first busbar of the first distribution box via the first DC / DC converter; the second battery is connected to a first busbar of the second distribution box via a second DC / DC converter; the third battery is connected to a first busbar of the third distribution box via a third DC / DC converter; and the fourth battery is connected to a first busbar of the fourth distribution box via the fourth DC / DC converter. The first busbar of the first distribution box is also electrically connected to a second busbar of the third distribution box, and the first busbar of the third distribution box is also electrically connected to a second busbar of the first distribution box; and / or the first busbar of the second distribution box is also electrically connected to a second busbar of the fourth distribution box, and the first busbar of the fourth distribution box is also electrically connected to a second busbar of the second distribution box.

[0013] According to another embodiment, the first battery and the second battery are located on the left side of the aircraft, and the third battery and the fourth battery are located on the right side of the aircraft.

[0014] According to yet another embodiment, the first busbar is electrically connected to the second busbar via a circuit breaker.

[0015] According to yet another embodiment, the first busbar is electrically connected to one of the plurality of DC / DC converters via a contactor.

[0016] According to a second aspect of this disclosure, an airborne power system for an all-electric aircraft is provided, comprising: a plurality of batteries for providing power; a plurality of DC / DC converters, each of the plurality of DC / DC converters being electrically connected to a corresponding battery among the plurality of batteries to perform DC / DC conversion on the DC power output from that battery; a plurality of first DC busbars, each of the plurality of first DC busbars being electrically connected to the output of one of the plurality of DC / DC converters to receive the DC / DC converted DC power to supply power to a first DC load of the all-electric aircraft; and a plurality of DC / AC inverters, each of the plurality of DC / AC inverters being electrically connected to... The system includes: a plurality of first DC busbars connected to one of the plurality of first DC busbars to convert the output from the first DC busbars into three-phase AC power; a plurality of AC busbars, each of which is electrically connected to the output of one of the plurality of DC / AC inverters, thereby supplying power to the AC load of the all-electric aircraft; a plurality of transformer rectifiers, each of which is electrically connected to one of the plurality of AC busbars to convert the output from the AC busbars into DC power; and a plurality of second DC busbars, each of which is electrically connected to the output of one of the plurality of transformer rectifiers, thereby supplying power to the second DC load of the all-electric aircraft.

[0017] According to one embodiment, the system further includes a plurality of emergency batteries, each of which is electrically connected to one of a plurality of second DC busbars, wherein under normal conditions the plurality of second DC busbars are powered by the plurality of transformer rectifiers, and in emergency situations the plurality of second DC busbars are powered by the plurality of emergency batteries.

[0018] According to another embodiment, each of the plurality of first DC busbars is redundant with at least one of the other first DC busbars in the plurality of first DC busbars, and / or each of the plurality of AC busbars is redundant with at least one of the other AC busbars in the plurality of AC busbars, and / or each of the plurality of second DC busbars is redundant with at least one of the other second DC busbars in the plurality of second DC busbars.

[0019] According to another embodiment, each of the plurality of first DC busbars is electrically connected to at least one of the other first DC busbars in the plurality of first DC busbars via a contactor, and / or each of the plurality of AC busbars is electrically connected to at least one of the other AC busbars in the plurality of AC busbars via a contactor, and / or each of the plurality of second DC busbars is electrically connected to at least one of the other second DC busbars in the plurality of second DC busbars via a contactor.

[0020] According to a third aspect of this disclosure, a power system for an all-electric aircraft is provided, comprising: a power system according to a first aspect of this disclosure; and an airborne power system according to a second aspect of this disclosure, wherein the power system is electrically isolated from the airborne power system.

[0021] According to one embodiment, the power system and the airborne power system are placed in separate compartments to electrically isolate them.

[0022] According to a fourth aspect of this disclosure, an all-electric aircraft including the power system described in a third aspect of this disclosure is provided.

[0023] The aspects generally include, as substantially as described herein with reference to the accompanying drawings and as explained by the drawings, methods, apparatus, systems, computer program products and processing systems.

[0024] The foregoing has broadly outlined the features and technical advantages of the examples according to this disclosure so that the following detailed description may be better understood. Additional features and advantages will be described thereafter. The disclosed concepts and specific examples can be readily used as the basis for modifying or designing other structures for implementing the same purposes as this disclosure. Such equivalent constructions do not depart from the scope of the appended claims. The characteristics of the concepts disclosed herein, in both their organization and manner of operation, and their associated advantages, will be better understood by considering the following description in conjunction with the accompanying drawings. Each drawing is provided for illustrative and descriptive purposes and does not define any limitation on the claims. Attached Figure Description

[0025] To gain a more detailed understanding of the features described above in this disclosure, reference can be made to various aspects of the above-briefly summarized content, some of which are illustrated in the accompanying drawings. However, it should be noted that the drawings illustrate only certain typical aspects of this disclosure and should not be considered as limiting its scope, as other equivalent aspects are permissible in this description. Identical reference numerals in different drawings may identify the same or similar elements.

[0026] Figure 1 A schematic block diagram of a power system for an all-electric aircraft according to an embodiment of the present disclosure is shown;

[0027] Figure 2 A schematic block diagram of another power system for an all-electric aircraft according to an embodiment of the present disclosure is shown;

[0028] Figure 3 A schematic block diagram of an airborne electrical power system for an all-electric aircraft according to an embodiment of the present disclosure is shown;

[0029] Figure 4 A schematic block diagram of a power system for an all-electric aircraft according to an embodiment of the present disclosure is shown; and

[0030] Figure 5 A schematic diagram of an example all-electric aircraft according to an example embodiment of the present disclosure is shown. Detailed Implementation

[0031] The detailed description that follows, taken in conjunction with the accompanying drawings, is intended as a description of various configurations and is not intended to represent the only configurations in which the concepts described herein can be practiced. This detailed description includes specific details to provide a thorough understanding of the various concepts. However, it will be apparent to those skilled in the art that these concepts can be practiced without these specific details.

[0032] refer to Figure 1 The diagram illustrates a schematic block diagram of a power system 100 for an all-electric aircraft according to an embodiment of the present disclosure. It will be understood that the power system 100 can provide power to the aircraft.

[0033] like Figure 1 As shown, system 100 may include multiple batteries 1011, 1012 for providing power; multiple DC / DC converters 1031, 1032; multiple distribution panels 1051, 1052; and multiple DC / AC inverters 1071, 1072, 1073, 1074. Although in Figure 1 In this example, all components are shown as two, but it will be understood that this is merely an example and the number of components can be any suitable value.

[0034] In one embodiment, each of the plurality of DC / DC converters 1031, 1032 is electrically connected to a corresponding battery among the plurality of batteries 1011, 1012 to perform DC / DC conversion on the DC power output from that battery; each of the plurality of distribution boxes 1051, 1052 is electrically connected to the output of at least one of the plurality of DC / DC converters 1031, 1032 to receive the DC / DC converted DC power; each of the plurality of DC / AC inverters 1071, 1072, 1073, 1074 is electrically connected to the output of one of the plurality of distribution boxes 1051, 1052 to convert the DC output from the distribution box into three-phase AC power, thereby supplying power to the drive unit of the all-electric aircraft. Figure 1 The figure, indicated by reference numeral 110, includes left motors LM1 and LM2 and right motors RM1 and RM2, which provide electrical power.

[0035] In a preferred embodiment of this disclosure, the battery, DC / DC converter, power distribution box, and DC / AC converter included in the system 100 are symmetrically distributed along the central axis of the all-electric aircraft, and thus symmetrically distributed on the left and right sides of the all-electric aircraft.

[0036] In one embodiment, each of the plurality of batteries may include an independent battery management system (BMS). Figure 1 (Not shown in the image), and includes several high-power-density battery cells, battery casing, etc., which will not be described in detail here.

[0037] In one embodiment of this disclosure, DC / DC converters 1031 and 1032 are used to convert the DC power output from the battery into 270V high-voltage DC power. Distribution boxes 1051 and 1052 can be high-voltage DC distribution boxes, generally including contactors, control relays, circuit breakers, and busbars, used to control the safe and reliable operation of the power grid of the power system, wherein the busbars are electrically connected to the output of at least one of the multiple DC / DC converters 1031 and 1032. DC / AC inverters 1071, 1072, 1073, and 1074 can convert the 270V high-voltage DC power from DC / DC converters 1031 and 1032 via distribution boxes 1051 and 1052 into three-phase AC power to provide power to the propulsion motors of the all-electric aircraft.

[0038] In one embodiment, the propulsion motor / propeller propulsion system 110 is an electric propulsion power unit consisting of a three-phase AC motor driving a propeller. The number of propellers can be determined according to the overall aircraft architecture design, for example... Figure 1 Four are shown. The propulsion motor / propeller propulsion system 110 is generally installed symmetrically on the left and right sides of the aircraft, and the number of DC / AC inverters is equal to the number of propulsion motor / propeller propulsion systems.

[0039] In one embodiment of this disclosure, each of the plurality of distribution panel boxes 1051, 1052 further includes a power controller for the status of the equipment inside the distribution panel box, for example... Figure 1 The busbar power controllers 1111 and 1112 are shown.

[0040] In one embodiment, each of the plurality of distribution boxes 1051, 1052 is redundant with at least one of the other distribution boxes 1051, 1052, such that if one of the distribution boxes 1051, 1052 fails, the redundant distribution box takes over the operation. For example, system 100 may include four distribution boxes, and if one of the distribution boxes fails, one of the remaining three distribution boxes can be selected in any suitable manner to take over the function of the failed distribution box.

[0041] In yet another embodiment of this disclosure, such as Figure 1 As shown, the plurality of distribution panels may include a first distribution panel 1051 and a second distribution panel 1052, the plurality of batteries may include a first battery 1011 and a second battery 1012, and the plurality of DC / DC converters may include a first DC / DC converter 1031 and a second DC / DC converter 1032. In this embodiment, the first battery 1011 is connected to the first distribution panel 1051 via the first DC / DC converter 1031, and the second battery 1012 is connected to the second distribution panel 1052 via the second DC / DC converter 1032. Further according to this embodiment, as... Figure 1 As shown, the first distribution panel box 1051 and the second distribution panel box 1052 are electrically connected together via a contactor 113, so that the two can be redundant to each other.

[0042] In a specific example of this disclosure, the power system may include four drive motors / propeller propulsion devices, each drive motor connected to a high-voltage DC distribution box via a three-phase DC / AC inverter. The high-voltage DC distribution box consists of two similarly configured distribution boxes, connected by a high-power contactor to form a redundant power supply system. That is, if one distribution box fails to supply power, the other distribution box can supply power to the failed distribution box by closing the high-power contactor. Each distribution box is connected to two drive motors / propeller propulsion devices, and the status of the internal equipment is controlled by a busbar power controller. A high-power-density battery pack may be divided into two groups, left and right, each supplying power to the left and right distribution boxes via DC / DC converters.

[0043] In another embodiment of this disclosure, each of the plurality of distribution boxes may include a first busbar and a second busbar, wherein the first busbar is electrically connected between one of the plurality of DC / DC converters and the second busbar, and the second busbar is electrically connected between the first busbar and one of the plurality of DC / AC inverters. This is in Figure 2 This is shown in detail in the text.

[0044] The following is for reference. Figure 2 The diagram shows a schematic block diagram of another power system 200 for an all-electric aircraft according to an embodiment of the present disclosure.

[0045] like Figure 2 As shown, system 200 may include four batteries 2011, 2012, 2013, and 2014 for providing power; four DC / DC converters 2021, 2022, 2023, and 2024; four first busbars 2051, 2052, 2053, and 2054; four second busbars 2071, 2072, 2073, and 2074; and four DC / AC inverters 2091, 2092, 2093, and 2094. Figure 2 As can be seen, the first busbars 2051, 2052, 2053, and 2054 are electrically connected between the DC / DC converters 2021, 2022, 2023, and 2024 and the second busbars 2071, 2072, 2073, and 2074, respectively. Furthermore, the second busbars 2071, 2072, 2073, and 2074 are electrically connected between the first busbars 2051, 2052, 2053, and 2054 and the DC / AC inverters 2091, 2092, 2093, and 2094, respectively.

[0046] It is understood that the first busbar 2051 and the second busbar 2071 are included in the first distribution panel box ( Figure 2 In the second distribution panel box (not shown), the first busbar 2052 and the second busbar 2072 are included. Figure 2 (not shown in the image), and so on. Therefore, from... Figure 2 As can be seen, batteries 2011, 2012, 2013, and 2014 are connected to the first busbars 2051, 2052, 2053, and 2054 via DC / DC converters 2021, 2022, 2023, and 2024, respectively.

[0047] In a preferred embodiment of this disclosure, such as Figure 2 As shown, the first busbar 2051 of the first distribution panel box is also electrically connected to the second busbar 2073 of the third distribution panel box, and the first busbar 2053 of the third distribution panel box is also electrically connected to the second busbar 2071 of the first distribution panel box. In a further embodiment, as... Figure 2As shown, the first busbar 2052 of the second distribution box is also electrically connected to the second busbar 2074 of the fourth distribution box, and the first busbar 2054 of the fourth distribution box is also electrically connected to the second busbar 2072 of the second distribution box.

[0048] Those skilled in the art will understand that batteries 2011, 2012, 2013, and 2014, DC / DC converters 2021, 2022, 2023, and 2024, and DC / AC inverters 2091, 2092, 2093, and 2094 can be respectively connected to... Figure 1 The batteries 1011 and 1012, DC / DC converters 1031 and 1032, and DC / AC inverters 1071, 1072, 1073, and 1074 are similarly connected and operated, and will not be described in detail here.

[0049] In a preferred embodiment of this disclosure, the first battery 2011 and the second battery 2012 are located on the left side of the aircraft, and the third battery 2013 and the fourth battery 2014 are located on the right side of the aircraft.

[0050] In yet another embodiment of this disclosure, the first busbar may be electrically connected to the second busbar via a circuit breaker. For example... Figure 2 As shown, the first busbars 2051, 2052, 2053, and 2054 are electrically connected to the second busbars 2071, 2072, 2073, and 2074 via circuit breakers 2211, 2212, 2213, 2214, 2215, 2216, 2217, and 2218, respectively.

[0051] In yet another embodiment of this disclosure, the first busbar may be electrically connected via a contactor to one of a plurality of DC / DC converters. For example, from Figure 2 As can be seen, the first busbars 2051, 2052, 2053, and 2054 are electrically connected to the DC / DC converters 2021, 2022, 2023, and 2024 via contactors 2111, 2112, 2113, and 2114, respectively.

[0052] The inventors recognized that electric propulsion units ( Figure 2 The number of batteries (represented by reference numeral 210 in the attached figure) depends primarily on the current electric motor technology's ability to meet propulsion power and flight speed requirements, while the number of high-power-density batteries depends on the aircraft's safety requirements for the system, as well as the structural and performance factors of battery integration and installation. Therefore, Figure 2A preferred embodiment is shown, in which there are four groups of high-power-density batteries and four groups of drive motor / propeller propulsion systems, with the batteries and propulsion systems having a corresponding power supply relationship. Specifically, batteries 2011 and 2013 are redundant, and batteries 2012 and 2014 are also redundant, forming a redundant power supply architecture. This architecture is a superior solution in terms of overall reliability, balance, and implementation complexity of the power supply system.

[0053] refer to Figure 3 The diagram illustrates a schematic block diagram of an airborne electrical power system 300 for an all-electric aircraft according to an embodiment of the present disclosure. It will be understood that the airborne electrical power system 300 provides electrical power to the aircraft's onboard equipment, rather than powering drive motors.

[0054] like Figure 3 As shown, system 300 may include multiple batteries 3011, 3012 for providing power; multiple DC / DC converters 3031, 3032; multiple first DC busbars 3051, 3052; multiple DC / AC inverters 3071, 3072; multiple AC busbars 3091, 3092; multiple transformer rectifiers (TRUs) 3111, 3113; and multiple second DC busbars 3131, 3132. Although in Figure 3 In this example, all components are shown as two, but it will be understood that this is merely an example and the number of components can be any suitable value.

[0055] In one embodiment, each of the plurality of DC / DC converters is electrically connected to a corresponding battery among the plurality of batteries to perform DC / DC conversion on the DC power output by that battery; each of the plurality of first DC busbars is electrically connected to the output of one of the plurality of DC / DC converters to receive the DC / DC converted DC power to supply power to a first DC load of the all-electric aircraft; each of the plurality of DC / AC inverters is electrically connected to one of the plurality of first DC busbars to convert the output from the first DC busbars into three-phase AC power; each of the plurality of AC busbars is electrically connected to the output of one of the plurality of DC / AC inverters, thereby supplying power to an AC load of the all-electric aircraft; each of the plurality of transformer rectifiers is electrically connected to one of the plurality of AC busbars to convert the output from the AC busbars into DC power; and each of the plurality of second DC busbars is electrically connected to the output of one of the plurality of transformer rectifiers, thereby supplying power to a second DC load of the all-electric aircraft.

[0056] like Figure 3As shown, DC / DC converters 3031 and 3032 are electrically connected to batteries 3011 and 3012, respectively, to perform DC / DC conversion on the DC power output from one of batteries 3011 and 3012. First DC busbars 3051 and 3052 are electrically connected to the output of one of the DC / DC converters 3031 and 3032, respectively, to receive the DC / DC converted DC power to supply the first DC load of the all-electric aircraft. Figure 3 The multi-electric load shown is powered. DC / AC inverters 3071 and 3072 are electrically connected to the first DC busbars 3051 and 3052, respectively, to convert the output from the first DC busbars into three-phase AC power. AC busbars 3091 and 3092 are electrically connected to the outputs of DC / AC inverters 3071 and 3072, respectively, thereby supplying AC loads (as shown in the diagram) of the all-electric aircraft. Figure 3 The AC load shown is powered. TRUs (transformer rectifiers) 3111 and 3113 are electrically connected to AC busbars 3091 and 3092, respectively, to convert the output from the AC busbars into DC. Second DC busbars 3131 and 3132 are electrically connected to the outputs of TRUs 3111 and 3113, respectively, thereby supplying power to the second DC load of the all-electric aircraft (shown in the diagram). Figure 3 The low-voltage DC load shown in the figure is powered by the power supply.

[0057] In another embodiment of this disclosure, system 300 may further include a plurality of emergency batteries, each of which is electrically connected to one of the plurality of second DC busbars, wherein under normal conditions the plurality of second DC busbars are powered by the plurality of transformer rectifiers, and in emergency situations the plurality of second DC busbars are powered by the plurality of emergency batteries. Figure 3 As shown, system 300 also includes emergency batteries 3151 and 3152. Emergency batteries 3151 and 3152 are electrically connected to the second DC busbars 3131 and 3132, respectively. Thus, under normal conditions (i.e., without any faults), the second DC busbars 3131 and 3132 are powered by transformer rectifiers 3111 and 3113, and in emergency conditions (i.e., in the event of a fault), the second DC busbars 3131 and 3132 are powered by emergency batteries 3151 and 3152.

[0058] In another embodiment of this disclosure, each of the plurality of first DC buses is redundant with at least one of the other first DC buses, and / or each of the plurality of AC buses is redundant with at least one of the other AC buses, and / or each of the plurality of second DC buses is redundant with at least one of the other second DC buses. For example, as Figure 3As shown, the first DC busbar 3051 is connected to a contactor ( Figure 3 (Not shown in the image) is electrically connected to the first DC busbar 3052, making them redundant; the AC busbar 3091 is connected to a contactor ( Figure 3 The second DC busbar 3131 is electrically connected to the AC busbar 3082 (not shown in the image) to provide redundancy; the second DC busbar 3131 is connected to the AC busbar 3082 via a contactor (…). Figure 3 (Not shown in the image) is electrically connected to the second DC busbar 3132, making them redundant.

[0059] In a specific example of this disclosure, such as Figure 3 As shown, a typical airborne electrical system consists of left and right power supply channels. The left and right power supply channels are symmetrical. Taking the left power supply channel as an example, it includes the main power supply battery and the corresponding battery management system (BMS). Figure 3 (Not shown in the image), a DC / DC converter is used to convert the DC power output from the battery into 270V high-voltage DC power; a 270V left DC busbar is used to connect 270V high-voltage DC multi-loads; a DC / AC inverter connected to the 270V left DC busbar is used to convert the 270V high-voltage DC power into 115V three-phase AC power; a 115V left AC busbar is used to connect 115V AC loads; a TRU is used to convert 115V three-phase AC power into 28V DC power; a 28V left DC busbar is used to connect 28V low-voltage DC loads; a set of left emergency batteries and the corresponding battery controller ( Figure 3 (Not shown in the image). Under normal power supply conditions, the 28V left DC busbar is powered by the TRU; if the 28V left DC busbar loses power, it is powered by the left emergency battery. Therefore, the 28V left DC busbar has two power supply redundancies. The same architecture and power supply redundancy apply to the right power supply channel.

[0060] In addition, taking the left power supply channel as an example, under normal power supply conditions, the 270V emergency DC busbar is powered by the 270V left DC busbar, and the 115V emergency AC busbar is powered by the 115V left AC busbar. When the main power supply channel malfunctions, the emergency power supply channel provides power redundancy. Taking the right main battery failure as an example, the contactor between the 270V emergency DC busbar and the 270V left DC busbar disconnects, the contactor between the 115V emergency AC busbar and the 115V left AC busbar disconnects, the emergency high-power-density battery pack starts working to provide power to the 270V emergency DC busbar, and simultaneously closes the contactor between the 270V emergency DC busbar and the 270V right DC busbar, as well as the contactor between the 115V emergency AC busbar and the 115V right AC busbar, to provide power to the equipment in the right channel.

[0061] The following is for reference. Figure 4The diagram shows a schematic block diagram of a power system 400 for an all-electric aircraft according to an embodiment of the present disclosure.

[0062] like Figure 4 As shown, the power supply system 400 may include a power energy system 402 (such as those combined with...) Figure 1-2 The aforementioned power systems 100 and 200) and airborne power system 405 (such as combined with Figure 3 The airborne electrical system 300 is described above. Among them, the power electrical system 402 provides power to the propulsion motor 403, while the airborne electrical system 405 provides electrical energy to the airborne system 406.

[0063] like Figure 4 As shown, the power system 402 is electrically isolated from the airborne power system 405, as indicated by the dashed line 408. Thus, this disclosure provides a discrete aircraft power system for an all-electric aircraft.

[0064] In one embodiment of this disclosure, the power system 402 and the airborne power system 405 are located in separate compartments to electrically isolate them.

[0065] refer to Figure 5 The illustration shows a schematic diagram of an example all-electric aircraft 500 according to an exemplary embodiment of the present disclosure. In one embodiment, the all-electric aircraft 500 may include, according to... Figure 4 The power supply system 400 is described above.

[0066] Those skilled in the art will understand that although the term "battery" is used in this disclosure, it can be replaced by "battery pack," "a set of batteries," "a set of battery modules," etc., which will not be elaborated further here.

[0067] Those skilled in the art will also understand that although specific voltage values ​​are described in this disclosure, any other suitable voltage values ​​may be used, which will not be elaborated here.

[0068] The above detailed description includes references to the accompanying drawings, which form part of the detailed description. The drawings illustrate specific embodiments that can be practiced by way of illustration. These embodiments are also referred to herein as “examples.” Such examples may include elements other than those shown or described. However, examples including the shown or described elements are also contemplated. Furthermore, examples of any combination or arrangement of those elements shown or described are contemplated, or with reference to specific examples (or one or more aspects thereof) shown or described herein, or with reference to other examples (or one or more aspects thereof) shown or described herein.

[0069] In the appended claims, the terms “comprising” and “including” are open-ended, meaning that a system, apparatus, article of manufacture, or process containing elements other than those listed after such terms in a claim is still considered to fall within the scope of that claim. Furthermore, in the appended claims, the terms “first,” “second,” and “third,” etc., are used merely as designations and are not intended to indicate a numerical order of their contents.

[0070] Furthermore, the order of operations described in this specification is exemplary. In alternative embodiments, the operations may be performed in a different order than that shown in the accompanying drawings, and the operations may be combined into a single operation or broken down into more operations.

[0071] The above description is intended to be illustrative and not restrictive. For example, the examples described above (or one or more aspects thereof) may be used in conjunction with other embodiments. Other embodiments may be used by those skilled in the art after reviewing the above description. The abstract allows the reader to quickly determine the nature of this technical disclosure. This abstract is submitted and it is understood that it is not intended to interpret or limit the scope or meaning of the claims. Furthermore, in the above detailed description, various features may be grouped together to make this disclosure flow smoothly. However, the claims may not state every feature disclosed herein, as embodiments may characterize a subset of said features. Furthermore, embodiments may include fewer features than those disclosed in a particular example. Therefore, the appended claims are thus incorporated into the detailed description, with each claim existing independently as a separate embodiment. The scope of the embodiments disclosed herein should be determined by reference to the full scope of the appended claims and equivalents of such claims.

Claims

1. A power system for an all-electric aircraft, comprising: Multiple batteries used to provide power; A plurality of DC / DC converters, each of which is electrically connected to a corresponding battery among the plurality of batteries, for performing DC / DC conversion on the direct current output by that battery; Multiple distribution panels, each of which is electrically connected to the output of at least one of the multiple DC / DC converters to receive DC-DC converted power; as well as Multiple DC / AC inverters, each electrically connected to the output of one of the multiple distribution boxes, to convert DC output from the distribution box into three-phase AC power. Each of the plurality of distribution panels includes at least a contactor, a control relay, a circuit breaker, and a busbar, wherein the busbar is electrically connected to the output of at least one of the plurality of DC / DC converters. Each of the plurality of distribution panels includes a first busbar and a second busbar, wherein the first busbar is electrically connected between one of the plurality of DC / DC converters and the second busbar, and the second busbar is electrically connected between the first busbar and one of the plurality of DC / AC inverters, thereby the first busbar supplies power to the drive unit of the all-electric aircraft only through the second busbar. The plurality of distribution boxes includes a first, second, third, and fourth distribution box; the plurality of batteries includes a first, second, third, and fourth battery; the plurality of DC / DC converters includes a first, second, third, and fourth DC / DC converter; the first battery is connected to a first busbar of the first distribution box via the first DC / DC converter; the second battery is connected to a first busbar of the second distribution box via the second DC / DC converter; the third battery is connected to a first busbar of the third distribution box via the third DC / DC converter; and the fourth battery is connected to a first busbar of the fourth distribution box via the fourth DC / DC converter. Wherein: The first busbar of the first distribution box is also electrically connected to the second busbar of the third distribution box, and the first busbar of the third distribution box is also electrically connected to the second busbar of the first distribution box; and / or The first busbar of the second distribution box is also electrically connected to the second busbar of the fourth distribution box, and the first busbar of the fourth distribution box is also electrically connected to the second busbar of the second distribution box. The first busbar is electrically connected to the second busbar via a circuit breaker. The power system is electrically isolated from the onboard power system of the all-electric aircraft.

2. The power system according to claim 1, characterized in that, Each of the plurality of distribution boxes also has a power controller for the status of the equipment inside the distribution box.

3. The power system according to claim 2, characterized in that, Each of the plurality of distribution boxes is redundant with at least one of the other distribution boxes, such that if one of the plurality of distribution boxes fails, the redundant distribution box takes over the operation.

4. The power system according to claim 1, characterized in that, The first battery and the second battery are located on the left side of the aircraft, and the third battery and the fourth battery are located on the right side of the aircraft.

5. The power system according to claim 1, characterized in that, The first busbar is electrically connected to one of the plurality of DC / DC converters via a contactor.

6. A power system for an all-electric aircraft, comprising: The power system according to any one of claims 1-5; as well as The airborne electrical system includes: Multiple batteries used to provide power; A plurality of DC / DC converters, each of which is electrically connected to a corresponding battery among the plurality of batteries, for performing DC / DC conversion on the direct current output by that battery; A plurality of first DC busbars, each of which is electrically connected to the output of one of the plurality of DC / DC converters to receive DC-DC converted DC power for use in supplying power only to a first DC load of the all-electric aircraft; Multiple DC / AC inverters, each of which is electrically connected to one of the multiple first DC busbars to convert the output from the first DC busbar into three-phase AC power; Multiple AC busbars, each of which is electrically connected to the output of one of the multiple DC / AC inverters, thereby being used to supply power only to the AC load of the all-electric aircraft; A plurality of transformer rectifiers, each of which is electrically connected to one of a plurality of AC busbars to convert the output from the AC busbars into DC; and Multiple second DC busbars, each electrically connected to the output of one of the multiple transformer rectifiers, are used to supply power only to the second DC load of the all-electric aircraft. The first DC load, the AC load, and the second DC load are not drive motors used to provide power to the all-electric aircraft. The power system is electrically isolated from the airborne power system.

7. The power supply system according to claim 6, characterized in that, It also includes a plurality of emergency batteries, each of which is electrically connected to one of the plurality of second DC busbars, wherein under normal conditions the plurality of second DC busbars are powered by the plurality of transformer rectifiers, and in emergency situations the plurality of second DC busbars are powered by the plurality of emergency batteries.

8. The power supply system according to claim 6, characterized in that, Each of the plurality of first DC busbars is redundant with at least one of the other first DC busbars in the plurality of first DC busbars, and / or Each of the plurality of AC busbars is redundant with at least one of the other AC busbars in the plurality of AC busbars, and / or Each of the plurality of second DC busbars is redundant with at least one of the other second DC busbars in the plurality of second DC busbars.

9. The power supply system according to claim 6, characterized in that, Each of the plurality of first DC busbars is electrically connected to at least one of the other first DC busbars via a contactor, and / or Each of the plurality of AC busbars is electrically connected to at least one of the other AC busbars via a contactor, and / or Each of the plurality of second DC busbars is electrically connected to at least one of the other second DC busbars via a contactor.

10. The power supply system according to claim 6, characterized in that, The power system and the airborne power system are located in separate compartments to ensure electrical isolation between them.

11. An all-electric aircraft comprising the power system of any one of claims 6-10.

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