ENERGY MANAGEMENT SYSTEM FOR AN AIRCRAFT HAVING A HYBRID POWER SOURCE INCLUDING AT LEAST ONE RECHARGEABLE POWER SOURCE AND ONE GENERATING SOURCE - Patent application

Abstract

A power management system for an aircraft with a hybrid power source comprises at least one rechargeable power source and a power generation source, a detector (200) configured, on the one hand, to determine status data indicative of the status of elements of the aircraft's power consuming electric circuits controlled by the power management system and, on the other hand, to determine power data relating to the instantaneous power requested by the aircraft and / or the state of charge of the rechargeable power source of the aircraft, and an automaton (210) configured to receive the power data from the detector (200) and to determine a control state of the power source, the automaton (210) comprising at least three states of the group comprising: * a buffer state, in which the instantaneous power requested is less than the capacity of the power generation source and is provided by the latter, * a charging state, in which the instantaneous power requested is less than the capacity of the power generation source and is provided entirely by the power generation source, the power generation source generating surplus power used to charge the rechargeable power source, and * a turbo state, in which the instantaneous power requested is greater than the capacity of the power generation source and the rechargeable power source provides the make-up necessary to achieve the instantaneous power requested. The system further includes an adapter (220) configured to receive the status data and determine a backup electrical configuration if the status data indicates a fault; a controller (230) configured to receive the status information from the automaton (210) and determine electrical control of the rechargeable power sources (50, 60, 80, 90) and the power generation sources (18, 20) based on the instantaneous power required; and a switch (240) configured to issue commands to switches of power consuming electrical circuits of the aircraft controlled by the power management system to implement a nominal electrical configuration or, upon receiving a backup electrical configuration of the adapter (220), to implement the backup electrical configuration.

Description

[Technical field]

[0001] The present invention relates to the field of aircraft, and more particularly to the field of energy management systems for aircraft having a hybrid power source including at least one rechargeable power source and one power generating source. [Background technology]

[0002] The aviation industry is currently facing many changes, partly linked to the evolution of environmental requirements and partly linked to the development of electric aircraft. In particular, the VTOL sector (which stands for Vertical Take-off and Landing) is particularly dynamic, as it offers very interesting prospects as a new means of transportation.

[0003] Although VTOLs are a very old field (developed since 1921), their electrification has given rise to many new proposed solutions as well as regulations. In particular, the latest regulations (see, for example, SC-VTOL-01 Special Conditions for Vertical Take-Off and Landing (VTOL) Aircraft, published on July 2, 2019) require redundancy of all systems related to the power sources through propulsion and flight, motors, and the entire electrical system to ensure the continuity of flight (called "continuation of safe flight and landing") as well as emergency landings after the occurrence of a failure. This is also called "one-fail-safe", i.e. "resistant to simple failures".

[0004] These regulations have greatly encouraged the development of solutions with redundancy, not only at the level of the rechargeable power sources, but also at the level of the power generation sources, but also at the level of the motors and rotors.

[0005] The consequence of this development is a very significant complication of the control systems. Conventionally, in an aircraft with a hybrid power source, three control systems can take place: a Flight Management System (so-called "Flight Management System" or FMS), which manages the flight parameters according to the external conditions and the flight control devices, in particular determining the instantaneous power required by the motors and / or rotors and controlling them accordingly, but optionally other parameters such as the orientation of flaps, stabilizers, ailerons, etc.; A battery management system (so-called "battery management system", or BMS) intended to optimize the management of electrical energy in a rechargeable power source; and an energy management system (so-called "Energy Management System" or EMS) to manage the power supply to consumers according to the requirements of the FMS (as can be done in the case of conventional thermal propulsion systems by a FADEC, which stands for "Full Function Digital Engine Control").

[0006] These three systems may coexist, but it is also possible that the BMS and EMS may be combined, for example if the BMS or EMS is highly simplified.

[0007] To address the challenge that redundancy management represents, namely optimizing power consumption, specific controllers have been developed that are implanted adjacent to the FMS, BMS and EMS. The addition of one-fail-safe management further complicates the situation.

[0008] Considering the reliability requirements of the aviation industry, these problems are traditionally managed by trees that exhaustively enumerate all possible cases for each problem (energy management, redundancy / fault management) and allow to reproduce this completeness for each problem, as is the case for example in CN109094790.

[0009] However, as progress progresses, the tree becomes more and more complex to generate and manage, becoming a source of risk. In addition, as soon as one element of the architecture changes, the tree becomes obsolete and must be restarted from zero. Other solutions call for managing according to flight phases, such as in FR 3 084 318 or US 2011 / 0178648. Again, these approaches are brief and hardly adaptable.

[0010] For this reason, to date, no system exists that allows for efficient management of electrical power in a hybrid aircraft while making it possible to control the one-fail-safe problem. [Prior art documents] [Patent documents]

[0011] [Patent Document 1] China Patent Application Publication No. 109094790 [Patent Document 2] French Patent Application Publication No. 3084318 [Patent Document 3] US Patent Application Publication No. 2011 / 0178648 Summary of the Invention

[0012] The present invention aims to improve the situation. To this end, an energy management system for an aircraft having a hybrid power source including at least one rechargeable power source and an electricity generating source is provided, the system comprising a detector configured to determine, on the one hand, status data representative of the state of the elements of the power consuming electric circuits of the aircraft controlled by the energy management system, and, on the other hand, power data relating to the instantaneous power required by the aircraft and / or the state of charge of the rechargeable power sources of the aircraft, and an automaton configured to receive the power data from the detector and to determine the control state of the power sources.

[0013] The automaton is * a buffer state in which the instantaneous power demanded is lower than the capacity of at least one generating source and is thereby supplied; * a charging state in which the instantaneous power demanded is lower than the capacity of at least one power generation source and is fully supplied by at least one power generation source, and one or more power generation sources generate surplus power that is used to recharge at least one rechargeable power source; *A turbo state in which the instantaneous power required is higher than the capacity of the at least one power generating source and the at least one rechargeable power source provides the make-up necessary to reach the instantaneous power required.

[0014] The system also includes an adapter configured to receive the status data and determine a backup electrical configuration if the status data indicates a fault; a controller configured to receive the status information of the automaton and determine electrical control of the at least one rechargeable power source and the at least one power generating source according to the required instantaneous power; and a switch configured to issue controls to switches of power consuming electrical circuits of the aircraft controlled by the energy management system to implement a nominal electrical configuration or, if a backup electrical configuration of the adapter is received, to implement the backup electrical configuration.

[0015] This energy management system is particularly advantageous because it allows the introduction of a functional and simplified management. In fact, the automaton / adapter separation makes it possible to manage the power and exploit the potential redundancies of the architecture while managing the one-fail-safe. Another advantage of the architecture of this energy management system is that it is "independent" of the architecture of the hybrid aircraft. In fact, by separating the energy management from the management of the configuration of the circuits to implement the power consumption, the energy management system makes it possible to manage any degree of electrical or mechanical redundancy of the aircraft, since it only has to change the settings of the adapters. This energy management system therefore has a large scalability and can be rapidly deployed on many hybrid aircraft with very different architectures.

[0016] According to various embodiments, the invention can have one or more of the following features. The controller is configured to use the rechargeable power source during a transition in which the power provided by the at least one power generating source increases to a requested instantaneous power when the automaton is in a buffer state; The automaton is initialized with a buffer state and the following transition rules: transitioning from a buffer state or a charging state to a turbo state if the instantaneous power required is higher than the capacity of at least one generating source; transitioning from a turbo state to a buffer state if the instantaneous power required is lower than the capacity of at least one generating source; transitioning from the buffer state to a charging state if the required instantaneous power is lower than the capacity of at least one generating source and one or more rechargeable sources indicate a state of charge lower than a full charge threshold; transitioning from the Charging State to a Buffered State if the required instantaneous power is lower than the capacity of at least one generating source and one or more rechargeable sources indicate a charging state equal to a full charge threshold; The automaton further includes a silent state in which at least one power source is deactivated; The automaton has the following transition rules: transitioning from a buffer state, a charging state, or a turbo state to a silent state if a silent mode control is received by the detector; transitioning from the silent state to the buffered state when the silent mode control is deactivated; The controller is configured to receive a backup electrical configuration of the adapter when a fault is detected and to control the shutdown of one or more elements of the aircraft's power consuming electrical circuits controlled by the energy management system in accordance with the backup electrical configuration; The switch is configured to receive a state of the automaton and responsively issue a control to a switch of a power consuming electrical circuit of the aircraft that is controlled by the energy management system.

[0017] The invention also relates to an energy management method for an aircraft having a hybrid power source including at least one rechargeable power source and an electricity generating source, characterized in that the method comprises the following operations: a) determining, on the one hand, status data indicative of the state of the elements of the power-consuming electrical circuits of the aircraft controlled by the energy management method, and, on the other hand, power data relating to the instantaneous power required by the aircraft and / or the state of charge of the rechargeable power sources of the aircraft, b) sending the power data of operation a) to an automaton configured to determine a control state of the power source, the automaton: * a buffer state in which the instantaneous power demanded is lower than the capacity of at least one generating source and is thereby supplied; * a charging state in which the instantaneous power demanded is lower than the capacity of at least one power generation source and is fully supplied by at least one power generation source, and one or more power generation sources generate surplus power that is used to recharge at least one rechargeable power source; * a turbo state in which the instantaneous power required is higher than the capacity of the at least one power generating source and the at least one rechargeable power source provides the necessary make-up to reach the instantaneous power required; c) determining the electrical control of the at least one rechargeable power source and the at least one power generation source according to the required instantaneous power and the state of the automaton determined in b); d) detecting the occurrence of a fault from the status data and determining a backup electrical configuration; e) issuing controls to switches of power consuming electrical circuits of the aircraft controlled by the energy management system to implement a nominal electrical configuration or, if a backup electrical configuration was determined in operation d), to implement the backup electrical configuration.

[0018] The invention also relates to a computer program comprising instructions for implementing the apparatus according to the invention or for performing the method according to the invention when said computer program is run on a computer, and to a data carrier on which this computer program is recorded.

[0019] Other characteristics and advantages of the invention will become better apparent from a reading of the following description, derived from examples given for illustrative and non-limiting purposes, and from the drawings in which: [Brief description of the drawings]

[0020] [Figure 1] FIG. 1 is a schematic diagram of the electrical configuration of a hybrid aircraft including an energy management system according to the present invention. [Diagram 2] FIG. 2 is a schematic diagram of the energy management system of FIG. 1. [Diagram 3] FIG. 3 is a schematic diagram of the states of the automaton of FIG. 2 and the transitions between them. [Figure 4] FIG. 2 is a diagram illustrating the sequence of control (from the aircraft pilot to the energy management system) and their respective effects. [Diagram 5] FIG. 2 shows a typical control loop of an aircraft implementing a device according to the invention; DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0021] The following drawings and description contain elements of a specific nature by their very nature, and therefore they can be used not only to better understand the invention, but also to contribute to its definition, if desired.

[0022] This specification may contain material that may be protected by copyright. The owner of the patent shall have no objection to anyone reproducing this patent document or its description exactly as it appears in the official records. All other rights are fully reserved.

[0023] As can be seen in FIG. 1 , an aircraft 2 according to the present invention includes an energy management system 4 according to the present invention, two horizontal drive units 6 and 8, four vertical drive units 10, 12, 14 and 16, and two power generation sources 18 and 20.

[0024] This type of aircraft is highly innovative and particularly suitable for demonstrating the possibilities of the energy management system 4. However, the aircraft may have a simpler construction, for example one single horizontal drive unit, one or two vertical drive units and one single power generation source. Further alternatively, the aircraft may be other than of the VTOL type, but of another type, for example a "conventional" hybrid aircraft.

[0025] In the example described herein, the horizontal drive unit 6 (respectively 8) comprises a DC-AC converter 22 (respectively 32), an electric motor 24 (respectively 34) and a propulsion unit 26 (respectively 36), for example of the propeller type. The propulsion unit 26 (respectively 36) is configured to enable the aircraft to move forward in a substantially horizontal direction. In the example described herein, the propulsion unit 26 (respectively 36) consumes 80 kW of power in flight regime.

[0026] The horizontal drive unit 6 (respectively 8) is connected at its input to a switch 28 (respectively 38) which enables this input to be connected to the output of a vertical drive unit 10 (respectively 14) or 12 (respectively 16), as will be described below.

[0027] Vertical drive units 10 (12, 14, 16, respectively) include rotors 42 (46, 72, 76, respectively) driven by motors 52 (56, 82, 86, respectively) and rotors 44 (48, 74, 78, respectively) driven by motors 54 (58, 84, 84, 88, respectively). Motors 52 and 54 are powered by respective DC-AC converters 62 and 64 (66 and 68, 92 and 94, 96 and 98, respectively). DC-AC converters 62 and 64 (66 and 68, 92 and 94, 96 and 98, respectively) are connected to an electrical bus of vertical drive units 10 (12, 14, 16, respectively) to which batteries 50 (60, 80, 90, respectively) are connected, as well as inputs connected to the power distribution bus of power generation source 18, and inputs connected to the power distribution bus of power generation source 20. Finally, the electric buses of each of the vertical drive units 10 and 12 (14 and 16 respectively) are connected to their respective outputs and connected to a switch 28 (38 respectively). As can be seen below, batteries 50, 60, 80 and 90 together deliver 600 kW when delivering 100% of their capacity.

[0028] In the example described herein, each power generation source 18 (respectively 20) includes, on the one hand, a turbogenerator 100 (respectively 102) and an AC-DC converter 104 (respectively 106). In the example described herein, each turbogenerator can provide 40 kW at 100% of its capacity. Alternatively, the power generation source can be composed of other DC or AC power generation sources followed by AC-DC or DC-DC converters. These sources can thus be based on turbogenerators powered by conventional fuels, biofuels, or synthetic fuels. Still alternatively, hydrogen-based power sources such as fuel cells can be used.

[0029] As can be seen in FIG. 2, the energy management system 4 is arranged to control the power generation sources 18 and 20 on the one hand, and the switches 28 and 38 on the other hand, as well as various protection elements not shown in FIG.

[0030] Analyzing Figure 1, it appears that all motors and electrical elements are duplicated, which allows to ensure one-fail-safe. In fact, there are two horizontal drive units and four vertical drive units themselves forming two subunits connected to the same horizontal drive unit, as well as two power generation sources.

[0031] Beyond this rather conventional duplication, an electrical bus specific to each vertical drive unit, as well as a power distribution bus specific to each power generation source, makes it possible to reach this objective.

[0032] In fact, as can be seen below, the particular structure of the aircraft in Figure 1 makes it possible to have a real hybridization of power sources, as opposed to existing solutions consisting of a mere juxtaposition. Thus, depending on the power requirements, both the battery and the power generation source can work in concert. Nevertheless, this architecture makes it possible to consider the battery as a pure "energy buffer". The battery is treated completely passively and does not require any software or hardware intelligence other than the basic intelligence required to operate the battery system alone, for example a BMS (Battery Management System) to activate protections and raise the state. This is completely contrary to all existing solutions in which an element is specifically designed to optimize the operation of the battery, acting as a control, or an element is provided to compensate for potential weaknesses of the battery, but exclusively alternating, i.e. the battery and this element cannot work simultaneously.

[0033] Figure 2 shows a schematic diagram of the energy management system 4 of Figure 1. As shown in this figure, the energy management system 4 includes a detector 200, an automaton 210, an adapter 220, a controller 230, and a switch 240.

[0034] The detector 200 is a system configured to receive various data of the aircraft 2, which it optionally processes and transmits, in whole or in part, to the automaton 210 on the one hand and to the adapter 220 on the other hand.

[0035] The data received by the detector 200 is therefore of two main types: On the one hand, status data indicating the state (load level, temperature, limits, operating state, fault state, etc.) of the elements of the aircraft's power consuming electrical circuits controlled by the energy management system 4; and On the other hand, power data relating to the instantaneous power required by the motors 52, 54, 56, 58, 82, 84, 86, 88 of the rotors 42, 44, 46, 48, 72, 74, 76, 78 and / or the motors 24 and 34 of the propulsion units 26 and 36, and / or the state of charge of the rechargeable power sources 50, 60, 80 and 90 of the aircraft.

[0036] The detector 200 therefore has an overall view of the functional state of elements related to power consumption, i.e. the presence or absence of faults on the one hand, as well as the presence of a flight phase of the aircraft 2, as well as the energy state of these elements, i.e. their instantaneous state, as well as the instantaneous power requirements related to the flight of the aircraft 2, which are determined in response to the control of the FMS.

[0037] Next, the expression "instantaneous power required" always refers to the power required by the motors 52, 54, 56, 58, 82, 84, 86, 88 of the rotors 42, 44, 46, 48, 72, 74, 76, 78 and / or the motors 24 and 34 of the propulsion units 26 and 36, unless a different definition is explicitly mentioned.

[0038] In the example described herein, the automaton 210 is a finite automaton, one embodiment of which is shown in Figure 3. As can be seen from this figure, the automaton 210 has four possible states: a so-called "buffer" state 300, in which the instantaneous power required is lower than the capacity of and can be supplied by at least one of the power generating sources 18 and 20; a so-called "charging" state 310, in which the instantaneous power required is lower than the capacity of the at least one power generation source and is entirely supplied by the at least one power generation source, and the at least one power generation source produces surplus power that is used to recharge the at least one rechargeable power source; a so-called "turbo" state 320, in which the instantaneous power required is higher than the capacity of at least one generating source and at least one rechargeable power source provides the necessary supplement to reach the instantaneous power required; and An optional so-called "silent" state 330 in which the power generation sources 18 and 20 are intentionally shut down to reduce sound emissions, which may also reduce pollutant emissions.

[0039] The automaton 210 presents transitions that aim to ensure the following: On the one hand, a minimum risk in the state transition decisions in order to limit the risk of failures associated with the automaton 210 (and thus the energy management system 4); On the other hand, charge batteries 50, 60, 80, 90 as frequently as possible.

[0040] Thus, the state transition is based on two variables. The instantaneous power required, Battery charge rates 50, 60, 80, 90.

[0041] As a matter of priority, the buffer or charging state is switched to the turbo state as soon as the instantaneous power required exceeds the power capacity of the power generating sources 18 and 20. Indeed, in this case it is essential to have the batteries 50, 60, 80, 90 and the power generating sources 18 and 20 operating simultaneously in order to provide sufficient power to carry out the flight control.

[0042] Secondarily, if the instantaneous power required is lower than the power capacity of the power generating sources 18 and 20, the priority is to recharge the batteries.

[0043] Thus, the automaton 210 can transition from the buffer state to the charging state or the turbo state, from the charging state to the turbo state, and from the charging state to the buffer state or from the turbo state to the buffer state. However, it cannot transition from the turbo state to the charging state, but must transition to the buffer state first.

[0044] In this way, each state transition is coordinated by a single change of state, either with respect to the required instantaneous power or with respect to the battery's state of charge. The reliability of the automaton 210 is therefore improved, since the risk of several consecutive transitions or non-transitions is minimal.

[0045] The automaton 210 also refreshes the state with a frequency of about 250 Hz, so that it can switch from the turbo state to the charging state almost instantly throughout the entire buffer state, or this frequency can be comprised, for example, between 50 Hz and 2 kHz.

[0046] Likewise, the silent state 330 is optional to the extent that it depends on manual activation by the pilot of the aircraft 2. The latter therefore sends a control to switch off the power generation sources 18 and 20, which transitions any state to the silent state. As in the case of the charging and turbo states, when this control is stopped, the automaton 210 transitions again to the buffer state.

[0047] The buffer state thus consists of the start-up state and the base state, in that it allows the operation of the automaton 210 to be more reliable. It is noted that the parameters that define the transitions can be modified. The transition from the buffer state to the charging state can thus be adjusted by the charging thresholds of the batteries 50, 60, 80, 90, which themselves can be modified according to the operational state and / or flight program of the aircraft 2. Similarly, the power capacities of the power generation sources 18 and 20 can be modified according to the operational state, functional state and / or flight program of the aircraft 2. For example, in case of a failure, it may be necessary to operate one or more of the power generation sources 18 and 20 at a capacity higher than their rated capacity, for example 120%. The thresholds for transitioning to the turbo state should be adapted accordingly. This adaptation is performed in the example described herein by the adapter 220, which adapts the transition thresholds, and by the controller 230, which transmits this information to the control board of the power generation source (respectively the control board of the power generation source).

[0048] Further alternatively, the switching to the silent state may not be purely manual, but may take into account the environment of the aircraft 2, for example taking into account one or more parameters of the flight height and the geographical location.

[0049] The adapter 220 is configured to receive status data from the detector 200. Based on this, the adapter 220 can determine a fault condition and a corresponding backup electrical configuration. For example, if the status data received from the detector 200 indicates that the motor 52 has failed, the adapter 220 will determine that the vertical drive unit 10 should be isolated and will return an electrical configuration indicating the need for this isolation, turning off the elements it contains.

[0050] Thus, adapter 220 includes a table of all possible fault configurations and a corresponding backup electrical configuration for each. Similarly, if adapter 220 determines that there is no fault, the adapter can return a nominal electrical configuration that can take into account, for example, the flight phase of aircraft 2. The fact that adapter 220 includes a table with all possible fault configurations is intended to compensate for any errors. Alternatively, adapter 220 can operate based on a logical operation to determine the backup electrical configuration.

[0051] In practice, the electrical configuration reflects the fact that isolation switches and / or turn-off controls should be activated whether in a nominal or backup configuration. In fact, each element of the circuit in Figure 1 is connected to the rest of the circuit by a switch (not shown), and the consumers and energy generators are further controlled as to when they are turned on or off.

[0052] Returning to the example of the failure of motor 52, adapter 220 determines a backup electrical configuration that isolates vertical drive unit 10 as shown in Figure 4. This results in a backup electrical configuration that indicates all of the switches in vertical drive unit 10, namely the switch between AC-DC converter 104 and the input of vertical drive unit 10 connected to DC-AC converter 62, the switch between AC-DC converter 106 and the input of vertical drive unit 10 connected to DC-AC converter 64, the switch between battery 50 and the rest of vertical drive unit 10, and the controls to turn off motors 52 and 54. This backup electrical configuration also indicates that all of the switches in vertical drive unit 12 should be activated to switch power to vertical drive unit 12, as well as switch 28 to switch to vertical drive unit 12.

[0053] The controller 230 receives, on the one hand, the state of the automaton 210 and, on the other hand, the electrical configuration of the adapter 220. Based on this, the controller 230 can control the rechargeable power sources 50, 60, 80 and 90 and / or the power generation sources 18 and 20 according to the energy regime corresponding to the state of the automaton 210 and the stop information indicated by the electrical configuration of the adapter 220.

[0054] for example, During the takeoff phase, in the activated silent state, the controller 230 sends a control to switch off the power generation sources 18 and 20. It should be understood that by switching off, the controller 230 issues a control for the electrical disconnection of the considered element of the electrical system. In practice, this control can result in a so-called super-idle regime, in which the considered power source rotates in a hyper-idle regime, avoiding switching off the power source for safety reasons (risk of restart problems).

[0055] Further alternatively, the shutdown may result in a literal electrical off. During the takeoff phase, in the charging state, the controller 230 sends controls to increase the power of the power generation sources 18 and 20 and reduce the power delivered by the batteries 50, 60, 80 and 90 to zero; During the turbo takeoff phase, the controller 230 sends controls to increase the power of the power generating sources 18 and 20 to their maximum capacity, During the cruise flight phase in the charged state, the controller 230 sends controls to increase the power of the power generating sources 18 and 20 until the power discharged by the batteries 50, 60, 80, and 90 is equal to the reciprocal of the charging power indicated by the BMS (i.e., in effect, the batteries receive this recharging power); During the cruise flight phase, in the buffer state, the controller 230 sends controls to increase the power of the power generating sources 18 and 20 in order to reduce the power delivered by the batteries 50, 60, 80 and 90 to zero; During the buffer state fall phase, the controller 230 sends controls to increase the power of the power generating sources 18 and 20 in order to reduce the power emitted by the batteries 50, 60, 80 and 90 to zero; During the landing phase in the buffer state, the controller 230 sends controls to increase the power of the power generation sources 18 and 20 in order to reduce the power delivered by the batteries 50, 60, 80, and 90 to zero; During the turbo landing phase, the controller 230 sends controls such as increasing the power of the power generating sources 18 and 20 to their maximum capacity.

[0056] Finally, the switch 240 is configured to perform an AND type binary operation between the configuration emitted by the adapter 220 on the one hand and the electrical configuration induced by the state of the automaton 210 on the other hand. Thus, returning to the example of a failure of the motor 52, while the state of the automaton 210 is the turbo state, for the switches of the vertical drive unit 10, this operation will be 0 (resulting from the separation ordered by the backup electrical configuration emitted by the adapter 220) times 1 (resulting from the turbo state) and back to 0 for these switches. Once the state of all the switches of the circuit has been determined by the switch 240, a corresponding control is sent to all the switches of the circuit, allowing the loop from the detector 200 to be restarted.

[0057] Therefore, it appears to be as follows: All elements of the circuit that consume or generate power are connected to the detector to indicate their state to the detector 200; All elements of the circuitry supplying power are controlled by a controller 230, All switches and on / off controls are controlled by switch 240 .

[0058] The energy management system of the present invention therefore completely separates the management of the power generation modes (via the automaton 210) from the management of the electrical configuration of the circuits for consuming this power (via the adapter 220). The controller 230 and the switch 240 are elements of an intentionally simplified nature, so that each is able to receive the outputs of the automaton 210 and the adapter 220 and take these outputs into account to control the power supply and the switch, respectively.

[0059] This de-correlation of the control of the electrical circuits is particularly innovative and advantageous in that it sets up an architecture that can be rapidly and reliably deployed on all types of hybrid aircraft, regardless of redundancy in their design.

[0060] Alternatively, the controller 230 may be configured to receive data only from the automaton 210 and the switch 240 may be configured to receive data only from the adapter 220 .

[0061] 5 shows a very high level operational flow chart of the architecture of an aircraft in which an energy management system according to the present invention is implemented. Thus, the control loop of the aircraft is continuous, with the FMS receiving control from the cockpit and pilot support elements via the function FMS() in operation 500.

[0062] Then, in an operation 510, the data generated by the function FMS() is propagated to the energy management system according to the invention via the function EMS() in which the system according to the invention described above is implemented. Finally, in an operation 520, the control issued by the controller 230 is implemented by the function DRV() which controls the individual elements. The loop then begins again with operation 500.

[0063] Of course, this loop is simplified compared to the actual operation of the aircraft, and interruptions may occur if a fault is detected or for other reasons that require some form of feedback. The purpose is to disclose at what level of the hardware architecture the invention sits; the FMS cannot be considered an EMS, nor can the reactor controller itself be considered an EMS.

Claims

1. An energy management system for an aircraft having a hybrid power source including at least one rechargeable power source (50, 60, 80, 90) and at least one power generation source (18, 20), comprising: a detector (200) configured to determine, on the one hand, status data indicative of the state of elements of power consuming electrical circuits of the aircraft controlled by the energy management system, and, on the other hand, power data relating to the instantaneous power required by the aircraft and / or the state of charge of the at least one rechargeable power source of the aircraft; An automaton (210) configured to receive the power data from the detector (200) and determine a control state of the hybrid power source, the automaton (210) comprising: a buffer state in which the instantaneous power required is lower than the capacity of the at least one power generation source, thereby providing; a charging state in which the instantaneous power required is lower than the capacity of the at least one power generation source and is fully supplied by the at least one power generation source, and one or more of the at least one power generation source generates surplus power that is used to recharge the at least one rechargeable power source; a turbo state in which the required instantaneous power is higher than the capacity of the at least one power generation source and the at least one rechargeable power source provides the necessary replenishment to reach the required instantaneous power; an adapter (220) configured to receive the status data and determine a backup electrical configuration if the status data indicates a fault; a controller (230) configured to receive state information of the automaton (210) and to determine an electrical control of the at least one rechargeable power source (50, 60, 80, 90) and the at least one power generation source (18, 20) according to the required instantaneous power; a switch (240) configured to issue controls to switches of the power consuming electrical circuits of the aircraft controlled by the energy management system to implement a nominal electrical configuration or, upon receipt of a backup electrical configuration of the adapter (220), to implement the backup electrical configuration; 2. An energy management system comprising: a controller configured to receive a backup electrical configuration of the adapter when a fault is detected; and to control a shutdown of one or more elements of the power consuming electrical circuits of the aircraft controlled by the energy management system in accordance with the backup electrical configuration.

2. 2. The energy management system of claim 1, wherein the controller (230) is configured to use the at least one rechargeable power source during a transition in which the power provided by the at least one power generation source increases to the required instantaneous power when the automaton (210) is in the buffer state.

3. The automaton (210) is initialized with the buffer state and has the following transition rules: transitioning from the buffer state or the charging state to the turbo state if the required instantaneous power is higher than the capacity of the at least one power generation source; transitioning from the turbo state to the buffer state if the required instantaneous power is lower than the capacity of the at least one power generation source; transitioning from the buffer state to the charging state if the required instantaneous power is lower than the capacity of the at least one power generation source and one or more rechargeable power sources have a state of charge lower than a full charge threshold; and transitioning from the charging state to the buffer state if the required instantaneous power is lower than the capacity of the at least one power generation source and one or more of the at least one rechargeable power sources indicates a state of charge equal to a full charge threshold.

4. The energy management system of claim 1 , wherein the automaton (210) further comprises a silent state in which the at least one power generation source (18, 20) is deactivated.

5. The automaton (210) has the following transition rules: transitioning from the buffer state, the charging state, or the turbo state to the silent state if a silent mode control is received by the detector (200); The energy management system of claim 4 , further comprising: transitioning from the silent state to the buffer state if the silent mode control is deactivated.

6. 6. The energy management system of claim 1, wherein the switch (240) is configured to receive a state of the automaton (210) and issue controls accordingly to the switches of the power consuming electrical circuits of the aircraft that are controlled by the energy management system.

7. An energy management method for an aircraft having a hybrid power source including at least one rechargeable power source (50, 60, 80, 90) and at least one power generation source (18, 20), comprising the steps of: a) determining, on the one hand, status data representative of the state of elements of the power consuming electrical circuits of the aircraft controlled by the energy management method, and, on the other hand, power data relating to the instantaneous power required by the aircraft and / or the state of charge of the rechargeable power sources of the aircraft; b) transmitting the power data of step a) to an automaton (210) configured to determine a control state of the hybrid power source, the automaton (210) a buffer state in which the instantaneous power required is lower than the capacity of the at least one power generation source, thereby providing; a charging state in which the required instantaneous power is lower than the capacity of the at least one power generation source and is fully supplied by the at least one power generation source, and one or more of the at least one power generation source generates surplus power that is used to recharge the at least one rechargeable power source; a turbo state in which the required instantaneous power is higher than the capacity of the at least one power generating source and the at least one rechargeable power source provides the necessary replenishment to reach the required instantaneous power; c) determining an electrical control of the at least one rechargeable power source (50, 60, 80, 90) and the at least one power generation source (18, 20) according to the required instantaneous power and the state of the automaton determined in step b); d) detecting the occurrence of a fault from the status data and determining a backup electrical configuration; and e) issuing controls to switches of power consuming electrical circuits of the aircraft controlled by an energy management system to implement a nominal electrical configuration or, if a backup electrical configuration was determined in step d), to implement this backup electrical configuration, and controlling shutdown of one or more elements of the power consuming electrical circuits of the aircraft controlled by the energy management system in accordance with the backup electrical configuration.

8. A computer program comprising instructions for performing the method according to claim 7 when said computer program is executed on a computer.

9. A data storage medium having the computer program according to claim 8 recorded thereon.

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