Electric propulsion system for an aircraft

Abstract

Electric Propulsion System for an Aircraft: An electric propulsion system (10) for an aircraft, comprising a plurality of branches (12), each branch (12) comprising: - at least one electrical power source (16), - at least one electric propulsion device (18), and - a branch bus (20) connected to each power source (16) and to each propulsion device (18) of the branch (12), characterized in that the propulsion system (10) further comprises a central bus (14), each branch (12) comprising a DC voltage converter (28) connecting the branch bus (20) to the central bus (14), said DC voltage converter (28) being configured to galvanically isolate said branch bus (20) from the central bus (14). Figure to be published with the abbreviation: 1

Description

Title of the invention: Electric propulsion system for an aircraft Technical field of the invention

[0001] The invention relates to an electric propulsion system for an aircraft. Such a system is notably installed on an aircraft with purely electric propulsion, for example an airplane. Prior art

[0002] Aircraft using propulsion that relies solely on electrical energy stored in batteries or other storage elements must include an electrical connection architecture capable of managing the energy flows from the storage elements to the electric motors regardless of the conditions of use and possible failures.

[0003] The electrical power distribution networks on board most current aircraft, intended for supplying secondary circuits (ventilation, lighting, heating, etc.), generally include an alternating network with alternating voltages of 115 V or 230 V for example, and / or a low voltage direct current network of 28 V. The electrical powers involved generally range from 50 to 500 kW.

[0004] An electrically powered aircraft requires greater electrical power and operates at higher DC voltages, on the order of 400 to 1000 V. For these electrically powered aircraft, it is necessary to define new electrical distribution architectures. Currently, two distribution architecture topologies are being studied: segregated architectures and distributed architectures.

[0005] Segregated architectures include several electrically disjoint parallel distribution channels, each comprising a power bus connecting a power source to one or more respective motors.

[0006] In such an architecture, the loss of an energy source inevitably leads to the loss of the associated engine(s). Depending on the type of aircraft and the type of propulsion system, the system may include a possibility of redirecting the engine(s) to a secondary backup energy source, of the same type as the primary source or of a different type (this is then referred to as a hybrid architecture).

[0007] However, such a redundant system results in increased bulk and on-board mass, by multiplying the connection buses and, where applicable, secondary sources that are not used in normal operation.

[0008] Distributed architectures seek to reduce this problem of clutter and mass, which is particularly critical for use on an aircraft.

[0009] In this type of architecture, the power sources are connected in parallel on a common power bus supplying all the motors. In nominal mode, the voltages between the different batteries balance so that the batteries share power to the different loads. In the event of a battery failure, the propulsion loads are powered by the remaining batteries.

[0010] This allows the batteries to be sized as close as possible to their nominal operating conditions, and thus optimizes battery mass. The electrical distribution network also becomes more integrated because the distribution bus associates several loads with several batteries.

[0011] A significant defect of distributed architectures is incident management, and in particular vulnerability to short circuits.

[0012] Indeed, in the event of a short circuit, a battery can deliver a current of several thousand amperes in a few milliseconds. When batteries are connected in parallel to the network, this phenomenon is multiplied by the number of batteries in parallel.

[0013] The first direct consequence is that the currents involved in this type of short circuit can be destructive to aeronautical components, and active protection solutions are not capable of isolating the fault.

[0014] The second consequence is related to the nature of the distributed distribution architecture. Indeed, unlike a segregated distribution architecture, which does not propagate the consequences of a short circuit to the entire propulsion network, in this type of architecture, the network voltage perceived by all the motors will collapse over a sufficiently long period to cause the unacceptable loss of propulsion of the aircraft, all the motors connected to the bus no longer being powered due to an automatic protection switch linked to a bus voltage level that is too low. Presentation of the invention

[0015] The invention aims to remedy these drawbacks by proposing an electrical propulsion network architecture that allows the resumption of propulsive loads in the event of the loss of one of the energy sources, without excessive bulk and mitigating the risks associated with a distributed architecture.

[0016] To this end, the invention relates to an electric propulsion system for an aircraft, the system comprising a plurality of branches, each branch comprising

[0017] - at least one source of electrical energy,

[0018] - at least one electric propulsion device, and

[0019] - a branch bus, connected to each source and each propulsion device of the branch,

[0020] characterized in that the propulsion system further comprises a central bus, each branch comprising a DC voltage converter connecting the branch bus to the central bus, said DC voltage converter being configured to galvanically isolate said branch bus from the central bus.

[0021] Such a system makes it possible to supply energy to a plurality of electric motors for the propulsion of an aircraft, allowing the motors of a branch to be restarted in the event of failure of the corresponding energy source, while being protected against the risks of propagation of short circuits.

[0022] The system may further include a backup source of electrical power connected to the central bus.

[0023] Such a feature makes it possible to compensate for the loss of an energy source in the event of failure and thus to maintain a sufficient voltage level in each propulsion device by distributing the voltage through the central bus.

[0024] Each propulsion device can be connected to one of the branch buses via a disconnecting element.

[0025] Such a feature allows individual protection of the propulsion device against overcurrents.

[0026] Each branch bus can be connected to the central bus by:

[0027] - a branch switching element of the contactor type or switch type semiconductor power, and

[0028] - a diode having a forward-biased direction oriented from the central bus towards the branch bus,

[0029] the branch switching unit and the diode being connected in series between the central bus and the branch bus, in parallel with the DC voltage converter.

[0030] Such a feature makes it possible to redistribute a portion of the electrical power in each branch to supply the propulsion devices of a branch whose source is failing, without risking back currents.

[0031] Each power source can be connected to the corresponding branch bus via a contactor-type source switching device or a semiconductor power switch-type device.

[0032] This feature makes it possible to isolate a failing source in order to restart the branch.

[0033] The system may include a control circuit configured to implement a failure management process in the event of a failure in the system.

[0034] Such a feature makes it possible to protect the propulsion system against various incidents while maintaining optimal capabilities according to the circumstances.

[0035] The invention also relates to a method for managing failures in an aircraft electric propulsion system as above, comprising steps of:

[0036] - prevention of fault propagation by means of galvanic isolation implemented by each DC voltage converter,

[0037] - localization of the fault and, if necessary, isolation of a source or a propulsion device, and

[0038] - where appropriate, pre-charging and then restarting stopped propulsion devices.

[0039] The method may include switching at least one source switching element and / or at least one branch switching element.

[0040] The method may include opening at least one disjunction member.

[0041] The invention also relates to an aircraft comprising an electric propulsion system as above. Brief description of the figures

[0042] [Fig-1] [Fig.1] is a schematic view of an electric propulsion system according to the invention,

[0043] [Fig.2][Fig.3][Fig.4] Figures 2 to 4 represent successive stages of a method for managing a failure in a source region of the propulsion system of the [Fig.1],

[0044] [Fig.5][Fig.6][Fig.7] Figures 5 to 7 represent successive stages of a method for managing a failure in a bus region of the propulsion system of the [Fig.1],

[0045] [Fig.8][Fig.9][Fig.10] Figures 8 to 10 represent successive steps in a process for managing a failure in a motor region of the propulsion system of the [Fig.1], and

[0046] [Fig. 11] [Fig. 11] represents a method for managing a failure in a central region of the propulsion system of [Fig. 1]. Detailed description of the invention

[0047] An electric propulsion system 10 according to the invention is shown in [Fig. 1]. The propulsion system 10 is intended to propel an aircraft independently, i.e. the aircraft is electrically powered only.

[0048] The propulsion system 10 comprises a plurality of branches 12, for example four branches 12 in the case shown, as well as a central bus 14.

[0049] Each branch includes an electrical power source 16, at least one electric propulsion device 18 and a branch bus 20 connected to said source 16 and to each propulsion device 18.

[0050] Sources 16 are, for example, batteries on board the aircraft, which supply the primary electrical network, intended for propulsion.

[0051] The propulsion devices 18 are, for example, electric motors installed in turbomachinery propelling the aircraft.

[0052] The central bus 14 and the branch buses 20 are electrical power distribution devices between several components, under high voltage and direct current, installed in secure enclosures. They also include logic connectors linked to a control circuit 21.

[0053] Each source 16 is connected to the branch bus 20 by a protected power line comprising a source switching element 22.

[0054] The source switching element 22 is, for example, of the contactor type or of the semiconductor power switch type. Such a switching element is suitable for switching even when carrying a relatively high electric current.

[0055] Each propulsion device 18 is connected to the corresponding branch bus 20 by a respective power line, including a disconnecting element 24, for example of the thermal fuse type. The disconnecting element 24 is adapted to passively disconnect the propulsion device 18 from the branch bus 20 in the event of an overcurrent flowing through it, or in a controlled manner by the control device 21.

[0056] Each branch bus 20 is connected to the central bus 14 via a DC voltage converter 26 and, arranged in parallel with said voltage converter 26, a diode 28 and a branch switching element 30.

[0057] The DC voltage converter 26 is adapted to galvanically isolate the branch bus 20 from the main bus 14.

[0058] Diode 28 is arranged with forward bias oriented from the central bus 14 to the corresponding branch bus 20. Diode 28 and branch switching element 30 together form a unidirectional electrical energy transfer circuit from the central bus 14 to the corresponding branch 12.

[0059] The branch switching element is open in normal operation of system 10.

[0060] The propulsion device 10 further includes an emergency power source 32 connected to the central bus 14 via an emergency switch 34.

[0061] The backup power source 32 is configured to supply power in the event of a failure of one of the sources 16, so as to compensate for the loss. The voltages in the various propulsion devices 18 are balanced through the central bus 14.

[0062] The control circuit 21 includes electronic boards and current and voltage sensors distributed throughout the propulsion system, in order to protect the entire network and to implement failure management processes described below.

[0063] For this purpose, four regions of the propulsion system 10 are defined in which a Failure can occur.

[0064] A source region 40 includes the power sources 16 and the entire power line upstream of the source switching unit 22.

[0065] A bus region 42 includes branch buses 20 and connecting lines to the central bus 14.

[0066] A motor region 44 includes the connection lines of the propulsion devices 18 to the associated branch bus 20.

[0067] Finally, a central region 46 includes the central bus 14, the connecting lines of this central bus 14 to the branch buses 20, and the DC voltage converters and associated distribution circuits.

[0068] A failure management process occurring in the source region 40 is described first. In Figures 2 to 11, the grey arrows denote a normal current corresponding to a standard operation of the system 10. The black arrows denote an abnormal current, in particular a short-circuit current, and the white arrows a correction current, put in place to compensate for faults in the system 10.

[0069] The fault is for example a short circuit, which is powered by the source 16 of branch 12 and the propulsion devices 18, which supply energy through their input capacities, as shown in [Fig.2].

[0070] Thanks to the galvanic isolation implemented by the DC voltage converter 26 of branch 12, only the source 16 of the branch supplies the short circuit, the other branches 12 being isolated.

[0071] In a first step, the control circuit 21 detects the fault using the current and voltage sensors of branch 12, and stops the regulation implemented by the DC voltage converter 26, in order to stop the distribution of energy in the propulsion devices 18 of branch 12.

[0072] In parallel, these propulsion devices 18 of branch 12 are also stopped for the time necessary to isolate the failure.

[0073] The operation of the other branches 12 continues unchanged, ensuring the propulsion of the aircraft.

[0074] During a second step, represented in [Fig.3], the source 16 of the branch 12 is isolated by switching the source switching element 22 associated by the control circuit 21.

[0075] Once the fault is isolated, the control circuit 21 restarts the regulation by the DC voltage converter 26 of branch 12, in order to allow a pre-charge of the propulsion devices 18.

[0076] Indeed, such a pre-charge is necessary before restarting the propulsion devices 18, whose voltage has dropped drastically.

[0077] Finally, once the pre-charge is complete, the branch switching element 30 is closed in order to transmit power from the DC voltage converters 26 of the other branches 12, to re-power the propulsion devices 18 of the failed branch, as shown in [Fig.4].

[0078] The additional power delivered by each other branch 12 to power the propulsion devices 18 of the failed branch is divided by the number of remaining branches, which reduces the individual load on each source 16.

[0079] Secondly, a failure management method taking place in the bus region 42, shown in [Fig.5], is described.

[0080] When an electrical fault occurs at the level of a branch bus 20, the first consequence resulting from this fault is the rise in the current delivered by the source 16.

[0081] As before, the sources 16 and the input capacities of the propulsion devices 18 deliver electrical energy in the short circuit.

[0082] Just as in the previous case, the galvanic isolation of the DC voltage converter 26 of branch 12 prevents the propagation of the fault to the other branches 12.

[0083] The control circuit 21 detects the presence of a short circuit in branch 12 using current and voltage sensors, and stops the regulation implemented by the DC voltage converter 26, in order to stop the distribution of energy in the propulsion devices 18 of branch 12.

[0084] During a second step, shown in [Fig.6], the control circuit 21 isolates the source 16 to stop the flow of current in the short circuit, by switching the source switching member 22. In parallel, the setpoint voltages of the propulsion devices 18 are set to zero.

[0085] In order to determine the location of the fault, the control circuit 21 initiates a pre-charge sequence via the DC voltage converter 26, as before. Since the fault is located on the branch bus 20, the pre-charge does not function and the voltage seen by the propulsion devices 18 does not increase sufficiently. During the pre-charge, the galvanic isolation implemented by the DC voltage converter 26 prevents any excessive current.

[0086] After a confirmation time, the control circuit determines that the fault is located on branch bus 20.

[0087] Branch 12 is then isolated, as shown in [Fig.7], because it is impossible to restart the associated propulsion devices 18.

[0088] Thirdly, a failure management method occurring in the motor region 44 is described.

[0089] In this case, the occurrence of a fault in the power supply line of one of the propulsion devices 18 results in an overcurrent from the source 16 and the still functional propulsion device 18, which drains its input capacity into the short circuit, as shown in [Fig.8].

[0090] As in previous cases, the galvanic isolation of the voltage converter 26 ensures that the short circuit does not propagate to the rest of the system 10.

[0091] The current and voltage sensors allow the control circuit 21 to discriminate the faulty propulsion device 18 and isolate it from the network, by triggering the opening of the corresponding disconnecting device 24, as shown in [Fig.9].

[0092] Once the disjunction member 24 is opened, the source 16 can again supply electrical power to the remaining propulsion device 18 of the hip 12, as shown [Fig. 10].

[0093] Once the voltage has returned to level, the control circuit 21 restarts the regulation of the DC voltage converter 26.

[0094] Finally, a failure management process taking place in the central region 46 is described.

[0095] In the event of an electrical fault occurring on the central bus 14, the DC voltage converters 26 cease their regulation. The control circuit 21 measures a significant voltage drop at the output of each DC voltage converter 26, thus saturating their regulation and switching to a malfunctioning mode, as shown [Fig. 11].

[0096] Once again, the galvanic isolation of each DC voltage converter 26 ensures that the fault does not propagate to the branches 12, preventing the sources 16 from delivering current into the fault.

[0097] In this case, stopping the DC voltage converters 26 is necessary and sufficient to isolate the fault.

[0098] The propulsion system 10 then operates as a conventional segregated architecture system, leaving enough energy and time for the aircraft to land safely.

[0099] In the four fault scenarios presented, the system retains its integrity and the propulsion devices remain functional, ensuring the aircraft's ability to perform its mission. The architecture of the propulsion system 10 also eliminates the need for short-circuit current protections sized for the number of onboard power sources 16. The presence of galvanically isolated DC voltage converters 26 could lead to a significant increase in mass, impacting the aircraft's flight capability. However, the architecture of the present invention minimizes this impact by allowing the converters to be sized not based on the effective power of each source, but rather on a reduced power rating, a function of the number of sources present.

Claims

Demands

1. An electric propulsion system (10) for an aircraft, comprising a plurality of branches (12), each branch (12) comprising: - at least one source (16) of electrical power, - at least one electric propulsion device (18), and - a branch bus (20), connected to each source (16) and to each propulsion device (18) of the branch (12), characterized in that the propulsion system (10) further comprises a central bus (14), each branch (12) comprising a DC voltage converter (28) connecting the branch bus (20) to the central bus (14), said DC voltage converter (28) being configured to galvanically isolate said branch bus (20) from the central bus (14).

2. System (10) according to claim 1, wherein the system (10) further comprises a backup source (32) of electrical power connected to the central bus (14).

3. System (10) according to any one of the preceding claims, wherein each propulsion device (18) is connected to one of the branch buses (20) via a disconnecting member (24).

4. System (10) according to any one of the preceding claims, wherein each branch bus (20) is connected to the central bus (14) by: - ​​a branch switching element (30) of the contactor type or of the semiconductor power switch type, and - a diode (28) having a forward direction oriented from the central bus (14) to the branch bus (20), the branch switching element (30) and the diode (28) being connected in series between the central bus (14) and the branch bus (20), in parallel with the DC voltage converter (26).

5. System (10) according to any one of the preceding claims, wherein each source (16) is connected to the corresponding branch bus (20) via a source switching device (22) of the contactor type or of the semiconductor power switch type.

6. System (10) according to the preceding claim, comprising a control circuit (31) configured to implement a failure management method in the event of a failure in the system (10).

7. A method for managing failures in an aircraft electric propulsion system (10) according to any one of the preceding claims, comprising the steps of:

8. - prevention of fault propagation by means of galvanic isolation implemented by each DC voltage converter (26), - localization of the fault and, if necessary, isolation of a source (16) or a propulsion device (18), and - where appropriate, pre-charge then restart of stopped propulsion devices (18). Aircraft comprising an electric propulsion system according to any one of claims 1 to 6.

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