Hybrid aircraft with transistorized power distribution protection connections.

Abstract

A hybrid aircraft is provided with a transistorized power distribution protection connection. The aircraft (2) includes drive units (6, 8, 10, 12, 14, 16), at least one battery (B1, ..., BM) for each drive unit (6, 8, 10, 12, 14, 16), at least one power generation source (26, 28) including at least one electric converter (E1, ..., EN) connected to each battery (B1, ..., BM), and a power supply control device (4) for controlling each power generation source (26, 28) according to the power requirements of the drive units (6, 8, 10, 12, 14, 16 and passively handling each battery (B1, ..., BM). Each electric converter (E1, ..., EN) is connected to each battery (B1, ..., BM) through a respective connection (102) having two metal-oxide semiconductor field-effect transistors (104, 106) connected in series with the forward direction reversed. Each connection (102) can operate according to only a unidirectional state, a bidirectional state, and a cut-off state in which current flows from the electric converter (E1, ..., EN) to the battery (B1, ..., BM).

Description

[Technical Field]

[0001] The field of the invention relates to aircraft, and more particularly to electric aircraft. [Background technology]

[0002] The aviation industry is currently undergoing numerous changes linked to environmental constraints, specifically the need to gradually reduce emissions of greenhouse gases such as carbon dioxide (CO2), so the development of electric aircraft is considered an important step.

[0003] A classic architecture for electric aircraft incorporates at least one power generation source configured to power a battery pack, which in turn provides power to the propulsion units as needed. For example, in the case of an electric vertical take-off and landing (eVTOL), the battery is combined with the vertical and horizontal propulsion units. Multiplying the battery count allows compliance with safety standards that, among other things, require component redundancy to ensure the continued safety of flight and landing.

[0004] In particular, in the case of a hybrid power aircraft (hybrid aircraft), the power generation source is a fuel-based electric energy generator such as a turbo engine or a fuel cell. Summary of the Invention [Problem to be solved by the invention]

[0005] A battery power supply circuit may be configured with batteries connected in parallel. Such a circuit is equipped with an isolation protection system that isolates the batteries in the event of a fault, particularly a short circuit. However, when such a protection system is activated, the batteries are isolated, resulting in a significant power loss.

[0006] One solution to prevent a failure from spreading to all batteries while suppressing power degradation is to directly separate the batteries or to simply separate them and provide a power supply circuit accordingly.

[0007] However, this separation principle does not work well in the case of hybrid aircraft. In fact, hybrid aircraft have the peculiarity that, for redundancy reasons, the number of batteries is generally greater than the number of power generation sources, and the batteries are connected to each other via a common power generation source. This interconnection point becomes a common point of failure between batteries powered by the same power generation source.

[0008] Moreover, each battery is usually connected to an electrical converter, such as an inverter, a rectifier, etc. As a result, if the number of batteries is increased several times, the number of electrical converters also increases, which has a significant impact on the weight of the aircraft, and therefore on the amount of electrical energy consumed during flight.

[0009] The present invention aims to improve the above situation. [Means for solving the problem]

[0010] Thus, the present invention provides a hybrid aircraft comprising: at least two drives, each including a propulsion unit and an electric motor; at least one source of stored electrical energy arranged to supply electrical energy to one or more of said electric motors; at least one power generation source including a fuel-based electric energy generator connected to a respective stored electric energy source; a power supply control device arranged to send power commands to said at least one power generation source depending on the power requirements of said drive; Equipped with the at least one stored electrical energy source is configured to provide electrical energy based on the power command in response to a difference between the power requirements of the drive unit and the power provided by the at least one power generation source; The at least one power generation source is further related to an aircraft, wherein the at least one stored electrical energy source is capable of recharging the at least one stored electrical energy source such that each stored electrical energy source is treated passively.

[0011] Each power generating source includes at least one electrical converter, and the at least one electrical converter converts, for each stored electrical energy source: a first metal oxide semiconductor field effect transistor and a second metal oxide semiconductor field effect transistor connected in series with opposite forward currents; or ● Two electrical circuits each including an insulated gate bipolar transistor and a parallel-connected reverse diode, connected in series with their forward directions facing in opposite directions; , and are connected via respective connections having the same reference numerals.

[0012] The power supply controller is configured to control the operation of each connection according to a set of states consisting of a unidirectional state in which current flows from the electrical converter to the source of stored electrical energy, a bidirectional state in which current flows in both directions, and a blocking state in which current flow is blocked.

[0013] In one or more embodiments, the first metal oxide semiconductor field effect transistor and the second metal oxide semiconductor field effect transistor in at least one connection are made of silicon carbide.

[0014] In one or more embodiments, at least one connection further comprises an electronic component, the electronic component comprising a resistor (120), the resistor being ● Metal oxide semiconductor field effect transistor, or An electric circuit having an insulated gate bipolar transistor and a parallel-connected reverse diode; and is connected in series with The electronic components are arranged to allow pre-charging of the corresponding stored electrical energy source.

[0015] In one or more embodiments, the at least one source of stored electrical energy is a battery.

[0016] In one or more embodiments, the drives include at least one takeoff drive and at least one cruise drive.

[0017] For example, at least one takeoff drive is a vertical takeoff and landing drive and at least one cruise drive is a horizontal drive.

[0018] In one or more embodiments, the fuel-based electrical energy generator of at least one power generation source is a turbo engine and the electrical converter of each of the power generation sources is an AC-DC converter.

[0019] The turbo engine may be powered by fuel, biofuel or synthetic fuel.

[0020] In one or more embodiments, the fuel-based electrical energy generator of at least one power generation source is a fuel cell and the electrical converter of each of the power generation sources is a DC-DC converter.

[0021] In one or more embodiments, the aircraft is configured to operate at least according to a turbo mode in which the power requirements of the drive require a supply of power from the at least one power generation source and the at least one stored electrical energy source, and the power supply controller controls each connection according to the unidirectional state.

[0022] In one or more embodiments, the aircraft is configured to operate at least according to an energy saving mode in which the power supply controller controls each connection according to the disconnection state.

[0023] In one or more embodiments, the aircraft is configured to operate at least according to a charge mode in which the power supply controller sends a power command to the at least one power generation source to charge the at least one stored source of electrical energy while meeting the power requirements of the drive unit.

[0024] Typically, in the charge mode, the power supply controller sequentially controls each connection through one or more charging phases, and the power supply controller is configured to implement each charging phase by associating a respective electrical converter with a stored electrical energy source and controlling each connection between the associated stored electrical energy source and the electrical converter in accordance with the unidirectional state and controlling any other connections in accordance with the interrupted state until each stored electrical energy source is charged.

[0025] Advantageously, the aircraft comprises a plurality of stored electrical energy sources, and the power supply controller in the charge mode is arranged to sequentially control each connection in one or more charging phases, the power supply controller being arranged to effect each charging phase by associating a respective electrical converter with a plurality of stored electrical energy sources and controlling each connection between the associated stored electrical energy source and the electrical converter in accordance with the bidirectional state and controlling any other connections in accordance with the interrupted state until each stored electrical energy source is charged.

[0026] In one or more embodiments, the aircraft is configured to operate at least according to a buffer mode in which the power supply controller sends power commands to the at least one power generation source to meet the power requirements of the drive unit through the at least one stored electrical energy source and the power supply controller sequentially controls each connection in one or more power supply phases, and the power supply controller is configured to implement each power supply phase by associating a respective electrical converter with a stored electrical energy source and controlling each connection between the associated stored electrical energy source and the electrical converter according to the unidirectional state and controlling all other connections according to the blocked state until the power requirements of the drive unit are met.

[0027] Other features, details and advantages will become apparent from a consideration of the following detailed description and from a review of the accompanying drawings. [Brief explanation of the drawings]

[0028] [Figure 1] 1 is a schematic diagram of the electrical architecture of an aircraft according to the present invention, with a single power generation source; [Figure 2] 1 is a schematic diagram of the electrical architecture of an aircraft according to the present invention, with two power generation sources; [Figure 3] 1 shows a circuit for powering a battery via a connection by an electrical converter in the electrical architecture of an aircraft according to the invention; FIG. [Figure 4] 10A and 10B are diagrams illustrating possible states of a connection portion. [Figure 5] FIG. 2 is a schematic diagram of the structure of a connection portion. [Figure 6] FIG. 6 is a schematic diagram showing one embodiment of the structure of FIG. 5. [Figure 7] 4 shows the power supply circuit of FIG. 3 in so-called "turbo mode" operation of the aircraft. [Figure 8] FIG. 8 is a diagram illustrating the turbo mode of FIG. 7 when the battery fails. [Figure 9]8 is a diagram illustrating the turbo mode of FIG. 7 when the electric converter fails. [Figure 10] 4 shows the power supply circuit of FIG. 3 in a so-called "energy saving mode" of operation of the aircraft. [Figure 11] 4 shows the power supply circuit of FIG. 3 in the so-called "charge mode" or in the so-called "buffer mode" of operation of the aircraft. [Figure 12] 12A and 12B illustrate the charge mode or buffer mode of FIG. 11 in special cases. DETAILED DESCRIPTION OF THE INVENTION

[0029] FIG. 1 shows an aircraft 2 equipped with a power supply control device 4, a number of drives 6, 8, 10, 12, 14, 16, a number of stored sources of electrical energy 18, 20, 22, 24, and a power generation source 26.

[0030] Typically, two of the drives 6, 8 are cruise drives utilized during the flight phase from takeoff to landing, and four of the drives 10, 12, 14, 16 are takeoff drives utilized during the takeoff phase.

[0031] As an example, the aircraft 2 may be an electric vertical take-off and landing (eVTOL) aircraft in which the four take-off drives 10, 12, 14, 16 are vertical drives and the two cruise drives 6, 8 are horizontal drives.

[0032] 1, cruise drive 6 includes a DC-to-AC converter 30, an electric motor 32, and a propulsion unit 34. Similarly, cruise drive 8 includes a DC-to-AC converter 38, an electric motor 40, and a propulsion unit 42.

[0033] The DC-AC converter 30 (or 38), which may also be referred to as an "inverter," is configured to generate alternating current from direct current.

[0034] The propulsion unit 34 (or 42) corresponds to, for example, a propeller or the like, and is provided to enable the aircraft 2 to move in a substantially horizontal direction. In flight mode, the propulsion unit 34 (or 42) consumes around 150 kilowatts (kW) of power.

[0035] Cruise drive 6 (or 8) is connected as an input to switching unit 36 ​​(or 44), which can connect the input to the output of takeoff drive 10 (or 14) or to the output of takeoff drive 12 (or 16).

[0036] The take-off drive 10 (or 12, 14 or 16) has a propeller 46 (or 50, 54 or 58) driven by an electric motor 62 (or 66, 70 or 74) and a propeller 48 (or 52, 56 or 60) driven by an electric motor 64 (or 68, 72 or 76).

[0037] In the context of the present invention, propellers 46, 48, 50, 52, 54, 56, 58, and 60 are also considered propulsion units, similar to propulsion units 34 and 42.

[0038] The electric motors 62, 64 (or 66, 68; 70, 72; or 74, 76) are powered by DC-AC converters 78, 80 (or 82, 84; 86, 88; or 90, 92), respectively, which are connected to the electrical bus of the takeoff drive 10 (or 12, 14, or 16).

[0039] The stored electrical energy source 18 (or 20, 22 or 24) is arranged to store electrical energy and supply it to the takeoff drive 10 (or 12, 14 or 16) in response to the power requirements of the takeoff drive 10 (or 12, 14 or 16). In addition, the stored electrical energy source 18, 20 (or 22, 24) is arranged to also supply electrical energy to the cruise drive 6 (or 8) via the switching unit 36 ​​(or 44).

[0040] For this purpose, the stored electrical energy source 18 (or 20, 22 or 24) is connected to a DC-AC converter 78, 80 (or 82, 84; 86, 88; or 90, 92) at the electrical bus of the take-off drive 10 (or 12, 14 or 16). Furthermore, the electrical bus of each take-off drive 10, 12 (or 14, 16) is connected to an output of the respective take-off drive 10, 12 (or 14, 16) and can be selectively connected to said output by the switching unit 36 ​​(or 44).

[0041] The stored electrical energy source 18 (or 20, 22, or 24) is typically a battery pack, i.e., a battery-to-battery capacitor, each intended to store electrical energy. Alternatively, the stored electrical energy source 18, 20, 22, or 24 may be a supercapacitor or a combination of a battery and a supercapacitor.

[0042] For the remainder of this specification, for ease of explanation, the source of stored electrical energy 18 (or 20, 22 or 24) will be referred to as the battery 18 (or 20, 22 or 24).

[0043] Typically, batteries 18, 20, 22, and 24 deliver a combined power of approximately 800 kilowatts (kW) when each is at 100% capacity.

[0044] A power generation source 26 is provided to generate and supply electrical energy to each of the batteries 18, 20, 22, 24. To this end, the power generation source 26 includes a number of power distribution buses.

[0045] 1, a power generation source 26 is connected to a first input of each of the takeoff drives 10, 12, 14, 16 by a first power distribution bus and to a second input by a second power distribution bus, which connect each of the takeoff drives 10, 12, 14, 16 with its corresponding battery 18, 20, 22, 24.

[0046] In the example of FIG. 1, the power generation source 26 includes two electrical converters 94 and 96 and a fuel-based electrical energy generator 98 .

[0047] More specifically, in this example, the electrical converters 94, 96 are AC-DC converters, and the fuel-to-electrical energy generator 98 is a turbo engine, such as a turbine generator, ie, a turbo generator.

[0048] An AC-DC converter 94 is connected to the first input of each takeoff drive 10, 12, 14, and 16. This means that the AC-DC converter 94 is the beginning of a first power distribution bus connecting the power generation source 26 to the first input of each takeoff drive 10, 12, 14, and 16. Similarly, an AC-DC converter 96 is connected to the second input of each takeoff drive 10, 12, 14, and 16. This means that the AC-DC converter 96 is the beginning of a second power distribution bus connecting the power generation source 26 to the second input of each takeoff drive 10, 12, 14, and 16.

[0049] The AC-DC converter 94 (or 96), which may be referred to as a "rectifier," is configured to generate direct current from alternating current.

[0050] Typically, turbo engine 98 is capable of delivering around 300 kilowatts (kW) of power at 100% capacity.

[0051] It should be noted that the power generating source 26 may operate on either direct current or alternating current, and the converters 94, 96 may be AC-DC converters or DC-DC converters as appropriate.

[0052] That is, power generation source 26 may be based on a turbo engine powered by a tank of conventional fuel, biofuel, or synthetic fuel (also known as synfuel). In such cases, electrical converters 94, 96 are AC-to-DC converters. Alternatively, power generation source 26 may be based on a hydrogen-based energy source, such as a fuel cell. In such cases, electrical converters 94, 96 are DC-to-DC converters. For purposes of this invention, such energy sources are considered fuel-based electrical energy generators.

[0053] The power supply control device 4 is a low-voltage device provided to control the power generating source 26, the switching units 36 and 44, and various other protective components not shown in FIG.

[0054] Rather than simply juxtaposing the batteries 18, 20, 22, 24, the electrical architecture of the aircraft 2 may provide for hybridization of the batteries 18, 20, 22, 24. That is, the batteries 18, 20, 22, 24 may cooperate with the power generation source 26 depending on the power requirements.

[0055] The batteries 18, 20, 22, and 24 are conventional batteries operating under a common control system (commonly known as a "battery management system" (BMS)). Such a system enables functions such as monitoring parameters such as voltage, temperature, charge rate, and state of health; preventing risks of deviations from expected operating ranges, such as overvoltage, overcurrent, and overheating; and optimizing battery performance. In the context of the present invention, no additional software or hardware intelligence is required. That is, the batteries 18, 20, 22, and 24 are treated passively in the sense that they require no special handling upon installation other than the connection of the batteries 18, 20, 22, and 24 to the electrical converters 94 and 96, as will be described later. From the rest of the electrical architecture of the aircraft 2, the batteries 18, 20, 22, and 24 simply appear as energy buffers. This contrasts with existing solutions, which have either a central component specifically designed to optimize battery operation or a component designed solely to compensate for possible battery failures (i.e., unable to operate simultaneously with the battery).

[0056] 1, the aircraft 2 is equipped with a single power generation source, namely, power generation source 26. However, it should be understood that in this example, the aircraft 2 may be equipped with multiple power generation sources.

[0057] 2 illustrates an embodiment in which the aircraft 2 is equipped with two power generation sources 26, 28. The power generation source 26 (or 28) includes an electrical converter 94 (or 96) and a fuel-based electrical generator 98 (or 100).

[0058] In the example of FIG. 2, the fuel-based electricity generator 98 (or 100) is a turbo engine and the electrical converter 94 (or 96) is an AC-to-DC converter.

[0059] Typically, each fuel-based electrical energy generator 98, 100 is capable of delivering approximately 150 kilowatts (kW) of electrical power at 100% capacity. Again, each power generating source 26, 28 may be based on a turbo engine powered by a tank of conventional fuel, biofuel, or synthetic fuel. Alternatively, a hydrogen tank-powered energy source such as a fuel cell may be used.

[0060] The electrical architecture of the aircraft 2 has been generally described with reference to FIGS.

[0061] As detailed above, the aircraft 2 includes at least one power generation source, such as the single power generation source 26 in FIG. 1 or the two power generation sources 26, 28 in FIG. 2, arranged to power one or more sources of stored electrical energy, such as the four batteries 18, 20, 22, 24 in FIGS. 1 and 2.

[0062] Because the aircraft 2 is a hybrid aircraft, it generally has more batteries than power generation sources. Furthermore, each distribution bus for each power generation source starts at an electrical converter (in this example, AC-DC converters 94 and 96), thereby reducing the number of electrical converters and reducing the weight of the aircraft 2. In other words, the electrical converters are located on the power generation source side rather than the battery side.

[0063] Illustratively, each electrical architecture shown in FIGS. 1 and 2 includes only two electrical converters 94, 96 for four batteries 18, 20, 22, 24.

[0064] However, the trade-off for this weight advantage for the aircraft 2 is that the batteries 18, 20, 22, 24 are interconnected by the power generating sources 26, 28. As a result, a short circuit occurring in the power generating sources or the batteries can spread.

[0065] To solve this problem, the applicant of the present application proposes a power supply circuit as shown in Figure 3. In the following description, we will consider how the electrical converter is connected to the battery.

[0066] 3 shows a circuit for feeding one or more stored electrical energy sources B1, ..., BM by one or more electrical converters E1, ..., EN, where M is a non-zero integer (natural number) corresponding to the number of stored electrical energy sources and N is a non-zero integer (natural number) corresponding to the number of electrical converters.

[0067] The power supply circuit described in this specification refers to the entire portion of the electrical architecture in Fig. 1 or 2 that relates to the electric converters 94 and 96 and the batteries 18, 20, 22, and 24. Therefore, if M=4 and N=2, the stored electric energy sources B1, B2, B3, and B4 correspond to the batteries 18, 20, 22, and 24, respectively, and the electric converters E1 and E2 correspond to the electric converters 94 and 96, respectively, as in Fig. 1 or 2.

[0068] For ease of explanation, the stored electrical energy sources B1, ..., BM will be referred to as batteries B1, ..., BM, respectively. It is noted that it is possible for there to be only one electrical converter (N=1) in the electrical architecture of the aircraft 2. Nevertheless, for the remainder of this specification, we will maintain generality and assume that there are multiple electrical converters E1, ..., EN and multiple batteries B1, ..., BM.

[0069] 3, each electric converter E1, ..., EN is connected to each battery B1, ..., BM by a respective connection 102. Consequently, the power supply circuit has as many connection parts as there are pairs of electric converters E1, ..., EN and batteries B1, ..., BM, i.e., N x M connection parts 102.

[0070] 4, the connection units 102 are configured to operate in only three states: a unidirectional state, a bidirectional state, and a cutoff state. More specifically, the operation of each connection unit 102 is controlled by the power supply control device 4.

[0071] In a unidirectional state, connection 102 allows current to flow from the electrical converter to the battery, and, of course, current cannot flow in the opposite direction, i.e., from the battery to the electrical converter.

[0072] In the bidirectional state, the connection 102 allows current to flow in both directions, ie, from the electrical converter to the battery, as well as from the battery to the electrical converter.

[0073] Finally, in the blocking state, connection 102 blocks current flow in either direction.

[0074] In this example, the connection unit 102 can only operate in these three states. Specifically, the power supply control device 4 cannot control the connection unit 102 to operate in a state where current can only flow from the battery to the electric converter.

[0075] The structure of the connection portion 102 will be described below with reference to FIGS.

[0076] In one embodiment shown in FIG. 5, the connection 102 is made in Metal Oxide Semiconductor Field Effect Transistor (better known as MOSFET) or Insulated Gate Field Effect Transistor technology.

[0077] More specifically, the connection 102 includes a first MOSFET 104 and a second MOSFET 106 connected in series. Each MOSFET 104, 106 includes three electrodes: a gate, a drain, and a source, which together form a semiconductor 108, 112. Each MOSFET 104, 106 also includes a diode 110, 114, which is a parasitic element formed by a pn junction between the drain and source. The diode 110, 114, sometimes referred to as a "body diode," is inherent in the MOSFET 104, 106.

[0078] The forward direction of each diode 110, 114 is opposite to the forward direction of the corresponding semiconductor 108, 112. In this example, the term "forward direction" refers to the direction of current flow. That is, in the semiconductors 108, 112, current flows from the drain to the source in the case of an N-type channel as shown in FIG. 5, and from the source to the drain in the case of a P-type channel.

[0079] 4, i.e., to enable operation in the unidirectional, bidirectional, and cutoff states, the first MOSFET 104 and the second MOSFET 106 are connected so that the forward directions of the respective semiconductors 108 and 112 are opposite to each other. Similarly, the forward directions of the diodes 110 and 114 of the first MOSFET 104 and the second MOSFET 106 are also opposite to each other. Therefore, the first MOSFET 104 and the second MOSFET 106 are connected in series so that the forward directions are opposite to each other.

[0080] Each of the MOSFETs 104 and 106 is, for example, a silicon carbide MOSFET, i.e., a SiC-MOSFET.

[0081] The connection 102 is in a unidirectional state when the first MOSFET 104 is in a blocking state and the second MOSFET 106 is in an on state. In other words, the connection 102 is in a unidirectional state when only the MOSFET whose forward direction is from the electrical converter to the battery is in an on state. The connection 102 is in a bidirectional state when the first MOSFET 104 and the second MOSFET 106 are in an on state. The connection 102 is in a cutoff state when the first MOSFET 104 and the second MOSFET 106 are in a blocking state.

[0082] Whether each MOSFET 104, 106 is in a blocking or on state is controlled by the power supply controller 4. To do this, the power supply controller 4 applies to the gate of each MOSFET 104, 106 the voltage required to switch it from a blocking state to an on state.

[0083] Alternatively, each MOSFET 104, 106 may be replaced by an insulated gate bipolar transistor (commonly known as an IGBT) connected in parallel with a reverse diode, i.e., a diode whose forward direction is opposite to that of the IGBT. In such an embodiment, each IGBT replaces the semiconductors 108, 112 of the MOSFETs 104, 106m, and each reverse diode replaces the diodes 110, 114.

[0084] In this case, the connection unit 102 has two electric circuits each including an IGBT and a reverse diode connected in parallel, connected in series. In these two electric circuits, the forward directions of the IGBTs and reverse diodes are opposite to each other when compared.

[0085] 6 shows an embodiment in which an electronic component 116 is connected to a first MOSFET 104 and a second MOSFET 106 within a connection portion 102. Specifically, the first MOSFET 104 and the second MOSFET 106 are connected in series with the positive wiring, while the electronic component 116 is disposed in the negative wiring.

[0086] The electronic component 116 includes a series-connected MOSFET 118 and a resistor 120. Similar to the first MOSFET 104 and the second MOSFET 106, the third MOSFET 118 includes a semiconductor 122 and a parasitic diode 124.

[0087] The electronic component 116 is arranged in the negative wiring so that the forward direction of the third MOSFET 118 is from the battery to the electrical converter. Equivalently, the forward direction of the diode 124 is from the electrical converter to the battery. In fact, as shown in FIG. 6, the third MOSFET 118 is oriented in the same direction as the first MOSFET 104.

[0088] The third MOSFET 118 is, for example, a SiC-MOSFET. Again, the third MOSFET 118 may be replaced with an IGBT connected in parallel with a reverse diode.

[0089] Electronic component 116 allows pre-charging of the corresponding battery via the negative line.

[0090] The use of transistors, whether MOSFETs or IGBTs, allows for a smaller mass of the connection 102 compared to solutions where the connection 102 is made of electromechanical contactors or relays. As an example, the mass of a SiC-MOSFET is about 6 g (grams) compared to about 750 g (grams) for electromechanical relays. In addition to being heavy and bulky, electromechanical contactors and relays also have the disadvantage of having a high risk of failure (friction, sparks) when powering on and off due to the presence of moving parts such as contacts.

[0091] As will be described in more detail below, the proposed power supply circuit, and in particular the use of connection 102, allows for both rated operation of aircraft 2 and failure situations (i.e., when at least one battery is unavailable or at least one electrical converter is unavailable).

[0092] Figure 7 shows one of the operating modes of the aircraft 2, i.e. turbo mode, when the power requirements of the drive units (more precisely, the respective electric motors) are so high that the batteries B1, ..., BM and the power generation sources (i.e., the electric converters E1, ..., EN) are utilized to their full capacity.

[0093] At this time, the power supply control device 4 commands each connection 102 to operate in a unidirectional state. This means that each electrical converter E1, ..., EN supplies each battery B1, ..., BM. If one of the batteries B1, ..., BM fails, for example due to a short circuit, the current generated by the short circuit is blocked by the connection 102 connected to the failed battery, so there is no possibility of it propagating from that battery to the other batteries B1, ..., BM. The same applies if a fault occurs in an electrical converter. The current generated by the short circuit cannot flow from the battery to the failed electrical converter.

[0094] In either case, the power supply controller 4 may then isolate the failed component by instructing the connection 102 connected to the failed component to switch from a unidirectional state to a disconnected state.

[0095] 8, in which battery B1 fails, the power supply control device 4 isolates battery B1 by switching all connections 102 connecting electrical converters E1, ..., EN to battery B1 from the unidirectional state to the disconnected state. Furthermore, since a battery (in this example, battery B1) is not powered, the electrical energy originally intended for that battery may be distributed to other batteries (in this example, batteries B2, ..., BM). It should be understood that this power supply circuit is flexible enough to implement dynamic power distribution, i.e., to supply electrical energy to batteries with higher requirements than other batteries.

[0096] In the case shown in Figure 9 where a failure occurs in the electric converter E1, the power supply control device 4 isolates the electric converter E1 by switching all of the connections 102 connecting the electric converter E1 to the batteries B1, ..., BM from a unidirectional state to a disconnected state.

[0097] FIG. 10 shows one of the operating modes of the aircraft 2 when power is not required from the power generation source, ie, the electrical converters E1, . . . , EN, ie, an energy saving mode.

[0098] At this time, the power supply control device 4 commands each connection unit 102 to operate in accordance with the interruption state. Even if any of the batteries B1, ..., BM or any of the electric converters E1, ..., EN fails due to, for example, a short circuit, the current generated by the short circuit is interrupted by each connection unit 102, and there is no possibility of the current propagating from the battery.

[0099] Figure 11 shows one mode of operation of aircraft 2 when the power requirements of the drive units (more precisely, each electric motor) are low enough that the electric converters E1, ..., EN can supply electrical energy to the drive units through batteries B1, ..., BM while also charging the batteries B1, ..., BM, i.e., charge mode.

[0100] Here, each of the electric converters E1, . . . , EN is associated with one of the batteries B1, .

[0101] At this time, the power supply control device 4 instructs each connection part 102 so that the connection part 102 between the electric converter and the battery associated with the electric converter is in a unidirectional state, and the other connection parts 102, i.e., each connection part 102 between the non-associated battery and the electric converter, is in a disconnected state.

[0102] After the N batteries associated with each electrical converter have been charged, all electrical converters E1,...,EN are assigned a new (next) battery to be charged, and so on. That is, up to N batteries B1,...,BM are charged sequentially in each charging phase, i.e., each iteration. Of course, if the number of batteries waiting to be charged is strictly less than the number of electrical converters, then some electrical converters E1,...,EN may not be associated with a new (next) battery.

[0103] That is, FIG. 11 shows one iteration in which electrical converter E1 is associated with battery B1 and electrical converter EN is associated with battery BM.

[0104] If M is divisible by N, then the number of iterations required to charge all batteries B1, ..., BM is M / N. On the other hand, if M is not divisible by N, then the number of iterations required to charge all batteries B1, ..., BM is [M / N] + 1 (where [...] is the floor function). In the final iteration, which charges a number of batteries equal to the remainder of the Euclidean division of M / N, all electrical converters can be used to fast charge the remaining batteries.

[0105] Which of the N batteries is selected for charging in each iteration may depend on the charge level of each of the batteries B1,...,BM, e.g., the battery with the lowest charge level may be given priority, or conversely, the battery with the highest charge level may be given priority.

[0106] FIG. 11 also shows another operating mode of the aircraft 2, i.e., a buffer mode, in which the power requirements of the drive units (more precisely, the respective electric motors) are low but the batteries B1, ..., BM do not need to be charged.

[0107] The batteries B1, ..., BM are treated passively like energy buffers, in other words the power provided by the electrical converters E1, ..., EN simply passes through the batteries B1, ..., BM to power the drive.

[0108] As in the charge mode, the power supply control device 4 performs one or more power supply phases, i.e., iterations. During each iteration, each electrical converter E1, ..., EN is associated with one of the batteries B1, ..., BM. The power supply control device 4 controls each connection 102 so that the connection 102 between the electrical converter and its associated battery is in a unidirectional state, and the other connections 102, i.e., the connections 102 between the electrical converter and its associated battery, are in a disconnected state. The buffer mode ends when the power requirements of the drive unit are met.

[0109] In the example of Figure 11, each electrical converter is associated with one battery in a sequential manner, however, it is also possible to associate multiple batteries with each electrical converter in each iteration.

[0110] 12, therefore, each of the electric converters E1, ..., EN is associated with P batteries, where P is an integer (natural number) greater than or equal to 2. In this case, the electric converter E1 is associated with the first P batteries, i.e., batteries B1, ..., BP, and the electric converter EN is associated with the last P batteries, i.e., batteries BM-P+1, ..., BM.

[0111] At this time, the power supply control device 4 controls each connection 102 so that the connection 102 between the electric converter and the battery associated with the electric converter is in a bidirectional state, and the other connection 102, i.e., the connection 102 between the non-associated battery and the electric converter, is in a cut-off state.

[0112] 12, the connections 102 between each battery B1, ..., BP and the electric converter E1 are in a bidirectional state. Similarly, the connections 102 between each battery BM-P+1, ..., BM and the electric converter EN are also in a bidirectional state. Meanwhile, the connections 102 between each battery B1, ..., BP and the electric converters other than E1 are in a cutoff state. Similarly, the connections 102 between each battery BM-P+1, ..., BM and the electric converters other than EN are also in a cutoff state.

[0113] In other words, when a battery is considered and is connected to N electrical converters E1, ..., EN via N connection parts 102, in the case of Fig. 12, these N connection parts 102 are instructed by the power supply control device 4 as follows: the connection part 102 between the associated electrical converter and the battery is in a bidirectional state, and the other N-1 connection parts 102 are in a cut-off state.

[0114] In the charging mode of the aircraft 2, the embodiment of FIG. 12 has the advantage of reducing charging times.

[0115] By making the connection 102 between the electrical converter and the P batteries associated with the electrical converter bidirectional, cross-flow is achieved between these P batteries, allowing them to be viewed as a single battery from the perspective of the associated electrical converter.

[0116] As a result of this bidirectionality, a fault, such as a short circuit, in one battery will propagate to the P-1 other batteries associated with the same electrical converter, but this effect is limited to the P batteries and not to the remaining batteries because the connectors 102 connecting them to the electrical converter associated with the faulty battery are in an interrupted state.

[0117] The configuration of the connection section 102 shown in FIG. 12 may be applied not only to the charge mode but also to the buffer mode.

Claims

1. A hybrid aircraft (2), comprising: at least two drives (6, 8, 10, 12, 14, 16) each including a propulsion unit (34, 42, 46, 48, 50, 52, 54, 56, 58, 60) and an electric motor (32, 40, 62, 64, 66, 68, 70, 72, 74, 76); at least one source of stored electrical energy (18, 20, 22, 24) arranged to supply electrical energy to one or more of said electric motors (32, 40, 62, 64, 66, 68, 70, 72, 74, 76); at least one power generation source (26, 28) including a fuel-based electric energy generator (98, 100) connected to a respective stored electric energy source (18, 20, 22, 24); a power supply control device (4) arranged to output power commands to said at least one power generation source (26, 28) depending on the power requirements of said drives (6, 8, 10, 12, 14, 16); Equipped with the at least one stored electrical energy source (18, 20, 22, 24) is configured to supply electrical energy based on the power command in accordance with a difference between the power requirements of the drive unit (6, 8, 10, 12, 14, 16) and the power supplied by the at least one power generation source (26, 28); The at least one power generation source (26, 28) is further capable of recharging the at least one stored electrical energy source (18, 20, 22, 24) such that each stored electrical energy source (18, 20, 22, 24) is treated passively. In the aircraft (2), Each power generating source (26, 28) includes at least one electrical converter (94, 96), and the at least one electrical converter (94, 96) converts, for each stored electrical energy source (18, 20, 22, 24): a first metal oxide semiconductor field effect transistor (104) and a second metal oxide semiconductor field effect transistor (106) connected in series with opposite forward currents; or Two electrical circuits each including an insulated gate bipolar transistor and a parallel-connected reverse diode, connected in series with their forward directions facing each other; and connected via respective connections (102) having the power supply control device (4) is configured to control the operation of each connection (102) according to a set of states consisting of a unidirectional state in which current flows from the electrical converter (94, 96) to the stored electrical energy source (18, 20, 22, 24), a bidirectional state in which current flows in both directions, and a blocking state in which current flow is blocked.

2. 2. The aircraft (2) of claim 1, wherein the first metal oxide semiconductor field effect transistor (104) and the second metal oxide semiconductor field effect transistor (106) in at least one connection portion (102) are made of silicon carbide.

3. 3. The aircraft (2) according to claim 1 or 2, wherein at least one connection (102) further comprises an electronic component (116), the electronic component (116) comprising a resistor (120), the resistor (120) being a metal oxide semiconductor field effect transistor (118), or An electric circuit having an insulated gate bipolar transistor and a parallel-connected reverse diode; and is connected in series with An aircraft (2), characterized in that said electronic components (116) are arranged to enable pre-charging of the corresponding stored electrical energy sources (18, 20, 22, 24).

4. 4. An aircraft (2) according to any one of claims 1 to 3, characterized in that at least one source of stored electrical energy (18, 20, 22, 24) is a battery.

5. 5. The aircraft (2) according to any one of claims 1 to 4, characterized in that the drives (6, 8, 10, 12, 14, 16) comprise at least one take-off drive (10, 12, 14, 16) and at least one cruise drive (6, 8).

6. 6. The aircraft (2) according to claim 5, characterized in that at least one take-off drive (10, 12, 14, 16) is a vertical take-off and landing drive and at least one cruise drive (6, 8) is a horizontal drive.

7. 7. The aircraft (2) according to any one of claims 1 to 6, characterized in that the fuel-based electrical energy generator (98, 100) of at least one power generation source (26, 28) is a turbo engine, and the respective electrical converter (94, 96) of the power generation source (26, 28) is an AC-DC converter.

8. 8. An aircraft (2) according to claim 7, characterized in that the turbo engine is powered by fuel, biofuel or synthetic fuel.

9. 9. The aircraft (2) according to any one of claims 1 to 8, characterized in that the fuel-based electrical energy generator (98, 100) of at least one power generation source (26, 28) is a fuel cell, and the respective electrical converter (94, 96) of the power generation source (26, 28) is a DC-DC converter.

10. 10. The aircraft (2) according to any one of claims 1 to 9, characterized in that the aircraft (2) is arranged to operate at least according to a turbo mode in which the power requirements of the drives (6, 8, 10, 12, 14, 16) require a power supply from the at least one power generation source (26, 28) and the at least one stored electrical energy source (18, 20, 22, 24) and the power supply control device (4) controls the respective connections (102) according to the unidirectional state.

11. 11. The aircraft (2) according to any one of claims 1 to 10, characterized in that the aircraft (2) is arranged to operate at least according to an energy saving mode in which the power supply control device (4) controls the respective connections (102) according to the disconnection state.

12. 12. The aircraft (2) according to any one of claims 1 to 11, characterized in that the aircraft (2) is arranged to operate at least according to a charge mode in which the power supply control device (4) outputs a power command to the at least one power generation source (26, 28) to charge the at least one stored electrical energy source (18, 20, 22, 24) while meeting the power requirements of the drive units (6, 8, 10, 12, 14, 16).

13. 13. The aircraft (2) according to claim 12, characterized in that the power supply control device (4) in the charge mode sequentially controls each connection (102) in one or more charging phases, and the power supply control device (4) is configured to implement each charging phase by associating a respective electric converter (94, 96) with a stored electric energy source (18, 20, 22, 24) and controlling each connection (102) between the associated stored electric energy source (18, 20, 22, 24) and the electric converter according to the unidirectional state and controlling any other connections (102) according to the interrupted state until each stored electric energy source (18, 20, 22, 24) is charged.

14. 13. The aircraft (2) according to claim 12, characterized in that the aircraft (2) comprises a plurality of stored electrical energy sources (18, 20, 22, 24), and the power supply control device (4) in the charge mode sequentially controls each of the connections (102) in one or more charging phases, and the power supply control device (4) is configured to perform each charging phase by associating a respective electric converter with each of a plurality of stored electrical energy sources (18, 20, 22, 24) and controlling each connection (102) between the associated stored electrical energy source (18, 20, 22, 24) and the electric converter in accordance with the bidirectional state and controlling all other connections (102) in accordance with the interrupted state until each stored electrical energy source (18, 20, 22, 24) is charged.

15. 15. The aircraft (2) according to any one of claims 1 to 14, wherein the aircraft (2) is adapted to operate at least according to a buffer mode in which the power supply control device (4) sends power commands to the at least one power generation source (26, 28) to satisfy the power requirements of the drives (6, 8, 10, 12, 14, 16) through the at least one stored electrical energy source (18, 20, 22, 24) and the power supply control device (4) sequentially controls each connection (102) in one or more power supply phases, and the supply control device (4) is configured to implement each power supply phase by associating a respective electric converter (94, 96) with each of the stored electric energy sources (18, 20, 22, 24) and controlling each connection (102) between the associated stored electric energy source (18, 20, 22, 24) and the electric converter according to the unidirectional state and controlling any other connections (102) according to the interrupted state until the power requirements of the drives (6, 8, 10, 12, 14, 16) are met.

Images