Fault-tolerant electric aircraft propulsion system

Abstract

Fault-tolerant aircraft electric propulsion system. Aircraft electric propulsion system comprising at least one electric motor (10) whose setpoint is calculated according to a first channel (18A) comprising a first control unit (20A) and according to a second channel (18B) comprising a second control unit (20B), the at least one electric motor comprising a rotor (12) and two star-wound stators (14A, 14B), and the first control unit of the first channel is connected to one of the two star-wound stators via a first power converter (22A), and the second control unit of the second channel is connected to the other of the two star-wound stators via a second power converter (22B), each of the first and second channels being sized at half the total power of the at least one electric motor. Figure for the abbreviation: Fig. 1.

Description

Title of the invention: Fault-tolerant electric aircraft propulsion system. Technical field

[0001] The present invention relates to the field of electric or hybrid aircraft propulsion and more particularly concerns an electric propulsion system architecture. Previous technique

[0002] It is well known that aeronautical regulations require aircraft propulsion systems to have high levels of reliability to meet safety requirements. Often, these levels of reliability translate into redundancies in certain components to compensate for their primary failure. This is particularly true for so-called ECS (Electronic Control System) electronics, which perform thrust regulation functions. Failure of these systems due to a single fault is only tolerated by aeronautical standards if it does not result in more than a 15% loss of performance.

[0003] Thus, in thermal propulsion architectures, a heat engine is associated with a control system called FADEC (Full Authority Digital Engine Control) which interfaces with the pilot's thrust commands and regulates engine thrust based on parameters transmitted by a series of sensors. In so-called Dual FADEC architectures, this critical engine control function is fully redundant to allow tolerance to the single failure of any component of the FADEC.

[0004] Similarly, in the context of new electric propulsion architectures, the electrical part includes not only a digital control unit, but also an inverter-type power converter and the motor's electrical components, consisting of its stator windings. To allow, as with a heat engine, simple fault tolerance for any part of the ECS, this ECS is also fully redundant.

[0005] Such redundancy is not without drawbacks, however, because it involves doubling the elements of the electrical power converter and the stator windings of the motor, which are bulky and heavy.

[0006] Furthermore, to ensure the resumption of regulation of the second ECS channel (the redundant channel) in the event of a failure of the first, it is necessary to provide a switching system to the motor of the contactor type, which, if mounted at the output of the power converter, must therefore be sized for high currents electrical components of the motor and no longer solely on the control signals as could be the case in Dual FADEC type thermal architectures.

[0007] This architectural choice therefore entails further mass and volume consequences compared to a simple chain to ensure the expected level of redundancy. Furthermore, this architectural choice with redundancy and a potential additional switching system complicates the system and reduces product availability due to a higher failure rate. Description of the invention

[0008] The main purpose of the present invention is therefore to overcome the aforementioned disadvantages by proposing an architecture that guarantees an ideal compromise between availability and reliability, and that makes it possible to ensure the security objectives while avoiding redundancy of bulky or massive elements.

[0009] These goals are achieved by an aircraft electric propulsion system comprising at least one electric motor whose setpoint is calculated according to a first channel comprising a first control unit and according to a second channel comprising a second control unit, characterized in that the at least one electric motor comprises a rotor and two star-wound stators and in that the first control unit of the first channel is connected to one of the two star-wound stators via a first power converter and the second control unit of the second channel is connected to the other of the two star-wound stators via a second power converter, each of the first and second channels being sized at half the total power of the at least one electric motor.

[0010] Thus, without having to double the power components, redundancies are limited to low-level electronic components, which have little impact on weight and onboard volume while allowing a good compromise between safety and availability. By taking advantage of intrinsic transient capabilities, this architecture makes it possible to compensate, for a limited time, for all or part of the power lost due to a simple failure.

[0011] Preferably, the first and second power converters are powered by two independent high-voltage DC power supplies, each sized at half the total power.

[0012] Advantageously, the first and second control units are powered from first and second low voltage DC power sources.

[0013] The aircraft propulsion system may further include diode OR logic to power the first and second control units from either of the first and second low voltage DC power sources.

[0014] Preferably, the wound stators have more than three phases on the same star or are wound on several three-phase stars with separate neutrals and in in which the number of lanes is equal to or greater than two, typically four or six.

[0015] The invention also relates to an aircraft electric motor regulation and monitoring system, comprising a regulation system configured to control at least one electric motor, the regulation system comprising: a first channel comprising a first control unit configured to calculate a setpoint for at least one electric motor based on measurement information from a measurement system, a second channel comprising a second control unit configured to calculate a setpoint for at least one electric motor based on measurement information from the measurement system,characterized in that at least one electric motor comprises a rotor and two star-wound stators, and in that the first control unit of the first channel is connected to one of the two star-wound stators via a first power converter, and the second control unit of the second channel is connected to the other of the two star-wound stators via a second power converter, each of the first and second channels being sized at half the total power of the electric motor.

[0016] Preferably, the control units have a dissimilar design to avoid any common failure.

[0017] According to the envisaged embodiment, the first power converter is connected to a first monitoring unit and the second power converter is connected to a second monitoring unit, or the first and second power converters are connected to a common monitoring unit.

[0018] Preferably, each monitoring unit includes a reset input in case of functional failure of these monitoring units.

[0019] Advantageously, the measurement system delivers the same data to the control units and monitoring units of the same track.

[0020] Preferably, the monitoring units are capable of stopping the power converters in the event of faults detected in this data. Brief description of the drawings

[0021] Other features and advantages of the present invention will become apparent from the description given below, with reference to the accompanying drawings which illustrate an example of an embodiment without being limiting in any way and on which:

[0022] [Fig-1] [Fig.1] illustrates an architecture of an electric propulsion system of an aircraft conforming to the invention,

[0023] [Fig.2] [Fig.2] shows a first example of the distribution of control and monitoring functions applied to the architecture of [Fig.1], and

[0024] [Fig.3] [Fig.3] shows a second example of the distribution of functions of control and monitoring applied to the architecture of [Fig. 1]. Description of the implementation methods

[0025] The principle of the invention is based on the possibility of constructing electric motors by dividing the stator into two star configurations, each supplying half the power to a single rotor. This half-power configuration therefore has the advantage not only of continuously maintaining half the power following a simple failure on either of the two paths, but also, by taking advantage of the transient capabilities of the electrical components, of allowing, thanks to the remaining path, the supply of higher transient power capable of compensating all or part of the power supplied at the time of the failure.

[0026] The architecture of the electric propulsion system is organized around an electric motor 10 comprising a rotor 12 and two star-wound stators 14A, 14B powered by an electronic control system (ECS 16) defining two separate channels 18A, 18B, each sized at half the total power of the electric motor. The first channel 18A comprises a first control unit 20A connected to one 14A of the two star-wound stators via a first power converter 22A, and the second channel 18B comprises a second control unit 20B connected to the other 14B of the two star-wound stators via a second power converter 22B.The two power converters (classically inverters) are each independently powered from a 24A, 24B high voltage DC power supply, each delivering half of the total power, and the two control units are each powered from a 26A, 26B low voltage DC power supply.

[0027] This particular configuration allows for a significant gain in mass and volume because the power converters have a power level divided by two. In addition, it allows for partial single-point fault tolerance with significant performance because not only does half of the total power of the electric motor remain available in continuous operation in the event of an internal failure on an electrical element of the propulsion chain (motor stator or ECS), but it can also have transient power on a single track equal to all or part of the total maximum motor power (when both tracks are supplied together), without oversizing, thanks to the intrinsic capabilities of permanent magnet synchronous electric machines to be able to operate at overspeed transiently.

[0028] This partial tolerance also applies to a simple failure of the aircraft's high-voltage power supply, since the two channels 18A, 18B are supplied with high-voltage direct current (HVDC) at half power by the two independent sources 24A, 24B. Thus, a failure of the power supply or the harness electric only results in the loss of half the power of the electric motor.

[0029] The invention further enables complete external fault tolerance of the DC low voltage power supplies 26A, 26B and of the control signals from the aircraft FADEC 28, due to the interfacing of the ECS, on two separate inputs, to these control signals and to the DC low voltage power supply.

[0030] Each interface therefore receives the same information from the aircraft, making it possible to overcome a failure of an aircraft component, for example the thrust lever or the control panel, or of the electrical connection harness. Internally, within the ECS, segregation management ensures that each channel of the ECS receives redundant control signals (digital and / or wired).

[0031] A diode logic 30 of the OR function type finally allows the interface to the low-voltage DC power supplies, ensuring fault tolerance of one of these two sources. It should be noted that the redundancy of these signals and interfaces is of little importance from the point of view of mass and volume.

[0032] Figures 2 and 3 illustrate in a very schematic way two examples of the aircraft engine regulation and monitoring device comprising a regulation system, an actuation system comprising at least one electric motor, and a measurement system (for example: temperature, speed, torque...).

[0033] The control system allows the electric motor 10 to be controlled according to a setpoint determined based on measurement information from the measurement system. This setpoint can be a position, speed, or torque setpoint, depending on the parameter of the controlled motor. It comprises two channels 18A and 18B, each configured to calculate a setpoint and control the electric motor accordingly. More generally, the different channels of the control system are therefore configured to determine one or more setpoints (depending on the number of electric motors involved).

[0034] On [Fig.2], the first channel and the second channel each include a COMA, COMB control unit configured to calculate a setpoint in a first way and to control the electric motor according to this setpoint via the power converter 22A, 22B.

[0035] Associated with either of the first and second channels, a MONA, MONB monitoring unit is, for example, configured to calculate a setpoint for the electric motor in a second way, different from the first way calculated by the COMA, COMB control unit. In particular, the calculation of the setpoint by the MONA, MONB monitoring unit is differentiated (both materially and functionally) compared to that implemented by the unit of COMA, COMB control. The MONA, MONB monitoring unit allows, in particular, verification that the COMA, COMB control unit is operating within its operational range (for example, a predetermined speed range). Indeed, this monitoring unit monitors the proper functioning of the control unit using the same data and has the authority to shut down the track by stopping (STOPA, STOPB) the power converter in the event of a detected fault (overspeed or excessive temperature) that is normally taken into account during normal operation.

[0036] Generally speaking, the monitoring unit is necessary to provide a means of switching off the control unit that is "independent" of it, in order to protect against events feared under aeronautical certification (Catastrophic or Hazardous). It is therefore an ultimate safety barrier in the event of a failure of the control unit.

[0037] Similarly, the two COMA and COMB control units can have a dissimilar design, that is, different in their components and / or embedded technologies, and the two stators can either share the same magnetic circuit (lamination pack) with a segregated geometric arrangement, or have two different magnetic circuits, or even lack a ferromagnetic core. This eliminates the risk of common design-related failures that could lead to the simultaneous loss of both power paths. This dissimilar architecture is particularly justified on aeronautical platforms with a large number of electric motors, such as new electric aircraft with distributed propulsion or vertical takeoff and landing (VTOL) aircraft.

[0038] It should be noted that in the proposed architecture, exchanges (not illustrated) exist between the two control units COMA and COMB and are necessary for its proper functioning. However, communication between the control units and the monitoring units can be optional, as data loss has no detrimental effect on performance.

[0039] With [Fig. 3], the architecture is optimized by combining the two monitoring units into a single common MON while remaining compatible with availability and safety objectives, in particular through the use of a reset means to restart the monitoring unit in case of a functional failure. It should be noted that the loss of this common MON does not result in a loss of performance but only the loss of a level of safety that is nevertheless acceptable until the end of the mission, for example.

[0040] The measurement system transmits information from sensors to the COMA, COMB control units and the MONA, MONB monitoring units of the control system. The measurement resources can be common (shared) or segregated (similar or diverse) and potentially spatially separated between the different units. different control and monitoring units of the regulation system.

[0041] It should be noted that although the invention has been described with regard to a two-way split architecture managing half the power, this number of channels is not a limitation and can, for example, be greater than two (typically four to six). It is also possible to consider a split into a larger number n of channels, allowing the power capacity to be maintained in the event of a power failure (nl / n). For example, assuming a power of 50 kW per channel, it is possible to control a 1 MW motor by connecting 20 channels in parallel.

[0042] Furthermore, each channel of the electric motor and its control can have more than 3 phases (the minimum number for a stator) on the same star connection. Thus, the stator can be multi-phase (typically 5, 6, or 7) and / or comprise several three-phase stars arranged in parallel (with separate neutrals). In this case, the power converter will have as many inverter arms as there are phases in the stator, and the number of control and monitoring units will, of course, be adjusted accordingly.

[0043] Similarly, another embodiment allows each track to be sized with a power capacity greater than Ptotal / n, at least for a transient period. For example, for a desired total power of 100 kW, dividing the power into two tracks results in a 50 kW power rating for each track. Considering an additional 50% transient power capacity for each track, i.e., 75 kW, a single failure leading to the loss of one track will still allow the motor to transiently access this 75 kW power, which is therefore 75% of the initial power. This operating mode can ensure that safety objectives are met in the event of a first failure during transient operation while limiting oversizing.

[0044] Thus, the invention relates to an electric propulsion system architecture offering an ideal compromise between availability and reliability while ensuring safety objectives by:

[0045] - external fault tolerance on low-level interfaces (power supply and digital communications), and

[0046] - partial tolerance for internal electrical faults while avoiding redundancy large or massive elements.

[0047] It allows the provision of sufficient transient power to ensure an ascent of the aircraft enabling it to gain a safe altitude, as well as the provision of sufficient continuous power to allow the aircraft to maintain its altitude and allow it to return to a landing area.

Claims

Demands

1. Aircraft electric propulsion system comprising at least one electric motor (10) whose setpoint is calculated both according to a first channel (18A) comprising a first control unit (20A) and according to a second channel (18B) comprising a second control unit (20B), characterized in that the at least one electric motor comprises a rotor (12) and two star-wound stators (14A, 14B) and in that the first control unit (20A) of the first channel (18A) is connected to one of the two star-wound stators via a first power converter (22A) and the second control unit (20B) of the second channel (18B) is connected to the other of the two star-wound stators via a second power converter (22B), each of the first and second channels being sized at half the total power of the at least one electric motor.

2. Aircraft electric propulsion system according to claim 1, wherein the first and second power converters (22A, 22B) are powered by two independent high-voltage DC power sources (24A, 24B), each sized at half the total power.

3. Aircraft electric propulsion system according to claim 1 or claim 2, wherein the first and second control units (20A, 20B) are powered from first and second low voltage DC power sources (26A, 26B).

4. Aircraft electric propulsion system according to claim 3, further comprising a diode OR logic (30) for powering the first and second control units from either of the first and second low voltage DC power sources.

5. Aircraft electric propulsion system according to claim 1 or claim 2, wherein the wound stators have more than three phases on the same star or are wound on several three-phase stars with separate neutrals and wherein the number of paths is equal to or greater than two, typically four or six.

6. Aircraft electric motor control and monitoring system, comprising a control system configured to control at least one electric motor (10), the control system comprising: a first channel (18A) comprising a first control unit (20A) configured to calculate a setpoint for the at least one electric motor based on measurement information from a measurement system, a second channel (18B) comprising a second control unit (20B) configured to calculate a setpoint for at least one electric motor based on measurement information from the measurement system, characterized in that the at least one electric motor comprises a rotor (12) and two star-wound stators (14A, 14B) and in that the first control unit (20A) of the first channel (18A) is connected to one of the two star-wound stators via a first power converter (22A) and the second control unit (20B) of the second channel (18B) is connected to the other of the two star-wound stators via a second power converter (22B), each of the first and second channels being sized at half the total power of the electric motor.

7. Aircraft electric motor regulation and monitoring system according to claim 6, wherein the control units have a dissimilar design to avoid any common failure.

8. Aircraft electric motor regulation and monitoring system according to claim 6 or claim 7, wherein the first power converter (INVERTERA) is connected to a first monitoring unit (MONA) and the second power converter (INVERTERB) is connected to a second monitoring unit (MONB).

9. Aircraft electric motor regulation and monitoring system according to claim 6 or claim 7, wherein the first and second power converters (INVERTERA, INVERTERB) are connected to a common monitoring unit (MON).

10. Aircraft electric motor regulation and monitoring system according to claim 8 or claim 9, wherein the monitoring units (MONA, MONB, MON) each have a reset input (RESET) in case of functional failure of these monitoring units.

11. Aircraft electric motor control and monitoring system according to any one of claims 6 to 10, wherein the measurement system delivers the same data to the control units and monitoring units of the same channel.

12. Aircraft electric motor regulation and monitoring system according to claim 11, wherein the monitoring units are capable of stopping (STOPA, STOPB) the power converters in case of failures detected in this data.

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