Fault tolerant aircraft electrical propulsion system

Abstract

Aircraft electrical propulsion system comprising at least one electric motor (10) of which a setpoint is calculated according to a first channel (ISA) 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 rated for half the total power of the at least one electric motor.

Description

[0001] Description

[0002] Title of the invention: Fault-tolerant electric aircraft propulsion system

[0003] Technical Field

[0004] The present invention relates to the field of electric or hybrid aircraft propulsion and more particularly concerns an electric propulsion system architecture.

[0005] Prior art

[0006] It is well known that aviation regulations require high levels of reliability for aircraft propulsion systems to meet safety requirements. These levels of reliability often result in redundancies of certain components to compensate for their primary failure. This is particularly the case for the so-called ECS (Electronic Control System) electronics that provide thrust regulation functions, the failure of which due to a single breakdown is only tolerated by aviation standards on the condition that it does not cause more than a 15% loss of performance.

[0007] Thus, in the context of thermal propulsion architectures, a thermal engine is associated with a regulation system called FADEC (Full Authority Digital Engine Control) which makes it possible to interface with the pilot's thrust control orders and to regulate the engine thrust from the parameters transmitted to it by a series of sensors. In so-called Dual FADEC architectures, this critical engine regulation function is fully redundant to allow tolerance to a single failure of any component of this FADEC.

[0008] Similarly, in the context of new electric propulsion architectures, the electrical part includes not only a digital control part, but also an inverter-type electrical power converter and the electrical part of the motor consisting of its stator windings. To allow, like a thermal engine, the tolerance to simple failure of any part of the ECS, this ECS is also fully redundant.

[0009] However, such redundancy is not without drawbacks because it involves doubling the elements of the electrical power converter and the stator windings of the motor, which are bulky and heavy.

[0010] 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 the high electrical currents of the motor and no longer solely for the control signals as could be the case in Dual FADEC type thermal architectures.

[0011] This choice of architecture therefore still involves mass and volume consequences compared to a simple chain to ensure the expected level of redundancy. In addition, this choice of architecture with redundancy and potential additional switching system complicates the system and deteriorates product availability due to a higher failure rate.

[0012] Statement of the invention

[0013] The main aim of the present invention is therefore to overcome the aforementioned drawbacks by proposing an architecture guaranteeing an ideal compromise between availability and reliability, and making it possible to ensure security objectives while avoiding the redundancy of large or massive elements.

[0014] These aims are achieved by an aircraft electric propulsion system comprising at least one electric motor, a setpoint of which 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.

[0015] Thus, without having to duplicate the power parts, redundancies are limited to low-level electronic parts which have little impact on the weight and on-board volume while allowing a good compromise on the level of safety and availability. By taking advantage of the intrinsic transient capacities, this architecture makes it possible to compensate for a limited time all or part of the power lost due to a simple failure.

[0016] Preferably, the first and second power converters are powered by two high voltage DC power sources independent of each other and each sized at half the total power.

[0017] Advantageously, the first and second control units are powered from first and second low voltage DC power sources.

[0018] The aircraft propulsion system may further include diode OR logic for powering the first and second control units from either of the first and second low voltage DC power sources.

[0019] 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 which the number of channels is equal to or greater than two, typically four or six.

[0020] The invention also relates to a system for regulating and monitoring an aircraft electric motor, 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 the at least one electric motor as a function of measurement information from a measurement system, a second channel comprising a second control unit configured to calculate a setpoint for the at least one electric motor as a function of measurement information from the measurement system,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 electric motor.,

[0021] Preferably, the control units have a dissimilar design to avoid common failure.

[0022] According to the embodiment envisaged, 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.

[0023] Preferably, the monitoring units each have a reset input in the event of a functional failure of these monitoring units.

[0024] Advantageously, the measurement system delivers the same data to the control units and monitoring units of the same channel.

[0025] Preferably, the monitoring units are capable of shutting down the power converters in the event of faults detected in these data.

[0026] Brief description of the drawings

[0027] Other characteristics and advantages of the present invention will emerge from the description given below, with reference to the appended drawings which illustrate an exemplary embodiment thereof without any limiting character and in which:

[0028] [Fig. 1] Figure 1 illustrates an architecture of an aircraft electric propulsion system according to the invention,

[0029] [Fig. 2] Figure 2 shows a first example of distribution of control and monitoring functions applied to the architecture of Figure 1, and [Fig. 3] Figure 3 shows a second example of distribution of control and monitoring functions applied to the architecture of Figure 1.

[0030] Description of the embodiments

[0031] The principle of the invention is based on the possibility of constructing electric motors by dividing the stator into two stars, each supplying half the power to a single rotor. This dimensioning at half the electrical power therefore has the advantage not only of continuously maintaining half the power following a simple failure on one or other of the two paths, but also, by taking advantage of the transient capacities of the electrical components, of allowing, thanks to the remaining path, the supply of a higher transient power capable of compensating all or part of the power supplied at the time of the failure.

[0032] 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 paths 18A, 18B each sized at half the total power of the electric motor. The first path 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 path 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 (conventionally inverters) are each powered independently from a 24A, 24B high voltage DC power source, each delivering half of the total power, and the two control units are each powered from a 26A, 26B low voltage DC power source.

[0033] This particular configuration allows a significant gain in mass and volume because the power converters have a power level divided by two. In addition, it allows partial tolerance to single failure 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 (engine stator or ECS), but it can also have transient power on a single channel equal to all or part of the total maximum motor power (when the two channels are supplied together), without oversizing, thanks to the intrinsic capacities of permanent magnet synchronous electric machines to be able to operate in overspeed transiently.

[0034] This partial tolerance also applies to a simple failure of the aircraft's high voltage power supply because the two channels 18A, 18B are supplied with high voltage direct current (HVDC) of half the power by the two independent sources 24A, 24B. Thus, the failure of the power source or the electrical harness again only results in the loss of half the power of the electric motor.

[0035] The invention further makes it possible to cover complete tolerance to external faults of the low voltage DC power sources 26A, 26B and the control signals coming from the FADEC 28 of the aircraft, due to the interfacing of the ECS, on two separate inputs, to these control signals and to the low voltage DC power supply.

[0036] Each interface therefore receives the same information from the aircraft, making it possible to avoid a failure of an aircraft component, for example the thrust lever or the control panel, or 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).

[0037] Finally, a 30 diode logic of the OR function type allows the interface to the low voltage DC power sources, ensuring fault tolerance of one of these two sources. It should be noted that the redundancy of these signals and these interfaces has little impact from a mass and volume point of view.

[0038] Figures 2 and 3 very schematically illustrate 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, etc.).

[0039] The control system makes it possible to control the electric motor 10 according to a setpoint determined according to measurement information from the measurement system. This setpoint may be a position, speed or torque setpoint depending on the parameter of the controlled motor. It comprises the two channels 18A, 18B, each configured to calculate a setpoint and control the electric motor according to this setpoint. 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 concerned).

[0040] In Figure 2, the first channel and the second channel each comprise a control unit COMA, COMB configured to, on the one hand, calculate a setpoint in a first manner and, on the other hand, control the electric motor according to this setpoint via the power converter 22A, 22B.

[0041] Associated with both the first channel and the second channel, a monitoring unit MONA, MONB is for example configured to calculate a setpoint for the electric motor in a second way, different from the first way calculated by the control unit COMA, COMB. In particular, the calculation of the setpoint by the monitoring unit MONA, MONB is diversified (both materially and functionally) compared to that implemented by the control unit COMA, COMB. The monitoring unit MONA, MONB makes it possible in particular to verify that the control unit COMA, COMB is operating within its operational domain (for example a predetermined speed interval).In fact, this monitoring unit monitors the correct operation of the control unit from the same data with authority to put the channel in "shutdown" by stopping (STOPA, STOPB) the power converter in the event of detection of failure (overspeed or excessive temperature) which are normally taken into account in normal operation.

[0042] Generally speaking, the monitoring unit is necessary to have a means of cut-off "independent" of the control unit, in order to protect against events feared in the sense of aeronautical certification (Catastrophic or Hazardous). It is therefore an ultimate safety barrier in the event of failure of the control unit.

[0043] Similarly, the two COMA and COMB control units can have a dissimilar design, i.e. different in terms of its components and / or embedded technologies, and the two stators can either share the same magnetic circuit (sheet metal pack) with a segregated geometric distribution, or have two different magnetic circuits, or not have 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 take-off aircraft (VTOL).

[0044] It can be noted that in the proposed architecture exchanges (not shown) exist between the two control units COMA, COMB and are necessary for its proper functioning. On the other hand, communication between the control units and the monitoring units can be optional, the loss of data having no detrimental effect on performance.

[0045] With Figure 3, the architecture is optimized by combining the two monitoring units into a single common MON while being compatible with availability and security objectives, in particular by the use of a reset means (RESET) to restart the monitoring unit in the event of a functional failure. It should be noted that the loss of this common MON unit does not result in a loss of performance but only the loss of a security level which is however acceptable until the end of the mission for example.

[0046] The measurement system transmits information from sensors to the COMA, COMB control units and MONA, MONB monitoring units of the regulation system. The measurement resources can be common (shared) or segregated (similar or diversified) and potentially spatially separated, between the different control and monitoring units of the regulation system. It should be noted that if the invention has been described with regard to an architecture shared in two channels managing half of the power, this number of channels cannot be limiting and can for example be greater than two (typically four to six). It is also possible to envisage a sharing in a larger number n of channels making it possible to maintain nl / n power capacity during an electrical failure. For example, assuming a power of 50kW per channel, it is possible to envisage the control of a 1MW motor with the paralleling of 20 channels.

[0047] In addition, each of the electric motor and its control channels can have more than 3 phases (minimum number for a stator) on the same star. Thus, the stator can be multi-phase (typically 5, 6 or 7) and / or have 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 stator phases and the number of control and monitoring units will of course be adapted accordingly.

[0048] Similarly, another embodiment allows the dimensioning of each channel with a power capacity greater than Ptotal / n, at least for a transient duration. For example, for a desired total power of 100kW, a split into two channels gives a dimensioning of each channel at 50kW. By considering a transient power capacity of 50% additional for each channel, i.e. 75kW, a simple failure leading to the loss of a channel will still allow the motor to have this power of 75kW transiently available, which is therefore 75% of the initial power. This operating mode can allow the safety objective to be met in the event of a first failure during transient operations while limiting oversizing.

[0049] Thus, the invention relates to an electric propulsion system architecture presenting an ideal compromise between availability and reliability while ensuring safety objectives by:

[0050] - tolerance to external faults on low-level interfaces (power supply and digital communications), and - partial tolerance to internal electrical faults while avoiding redundancy of large or massive elements.

[0051] It allows the supply of sufficient transient power to ensure a climb of the aircraft allowing it to gain a safe altitude as well as the supply of sufficient continuous power to allow the aircraft to maintain its altitude and allow it to return to a landing area.

Claims

Claims

1. An aircraft electric propulsion system comprising at least one electric motor (10) having a setpoint calculated both according to a first path (18A) comprising a first control unit (20A) and according to a second path (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 path (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 path (18B) is connected to the other of the two star-wound stators via a second power converter (22B), each of the first and second paths being sized at half the total power of the at least one electric motor.

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

3. An 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. An aircraft electric propulsion system according to claim 3, further comprising 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. An aircraft electric propulsion system according to claim 1 or claim 2, wherein the wound stators comprise more than three phases on a single star or are wound on several three-phase stars with separate neutrals and in which 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 as a function of measurement information from a measurement system, a second channel (18B) comprising a second control unit (20B) configured to calculate a setpoint for the at least one electric motor as a function of 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. An aircraft electric motor control and monitoring system according to claim 6, wherein the control units have a dissimilar design to avoid common failure.

8. An aircraft electric motor control 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. An aircraft electric motor control 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. An aircraft electric motor control and monitoring system according to claim 8 or claim 9, wherein the monitoring units (MONA, MONB, MON) each comprise a reset input (RESET) in the event of a 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 to the monitoring units of the same channel.

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

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