On-aircraft power supply system with redundancy and continuity of power supply in the event of a failure, and associated aircraft

Abstract

The invention relates to an on-aircraft power supply system (10), comprising: - a primary battery (40) connected to a rotary machine (20); - at least one secondary battery (42), each supplying an auxiliary propulsion system with power; - a DC-to-DC conversion device (45) comprising at least two separate DC-to-DC conversion modules (48), each being connected to a respective electric battery (40, 42), each module receiving one DC power from the battery (40, 42) and converting it into another DC power delivered as output; - a selection device (50) comprising at least one selection module (52) for selecting a respective voltage from among a plurality of voltages (U1, U2), each module being connected to at least two conversion modules (48) and selecting, according to a selection rule, one voltage from among the distinct voltages (U1, U2) output by the conversion modules (48), and then delivering the selected voltage to a bus (39).

Description

[0001] Electrical power system on board an aircraft, with redundancy and continuity of electrical power supply in the event of failure, associated aircraft

[0002] The present invention relates to an electrical power supply system on board an aircraft, the power supply system being configured to be connected to a rotating machine to receive input electrical energy and to deliver output electrical energy to at least one power supply bus of electrical equipment(s) of the aircraft.

[0003] The invention also relates to an aircraft, in particular a drone, comprising such an electrical power supply system.

[0004] The invention relates to the field of light aircraft, for example certified light drones, typically having a mass of less than 200 kg, as well as electric light aviation and electric urban taxis.

[0005] New regulations for certified light drones, such as the "Special Condition for Light Unmanned Aircraft Systems - High Risk" regulation issued by the European Union Aviation Safety Agency (EASA), impose safety constraints equivalent to those of transport aircraft, but with the need for weight, fuel consumption, volume and costs reduced by a factor of 100.

[0006] Existing aircraft electrical power systems typically make extensive use of electromechanical and electronic (static switch) systems, such as electromechanical relays and power converters, controlled by electronic computers to perform complex voltage regulation, conversion, distribution, and electrical reconfiguration functions in the event of a failure. These functions require bulky electronic equipment that is difficult and expensive to certify, and is incompatible with the aforementioned objectives of light aircraft.

[0007] Indeed, electromechanical relays are designed to cut off the supply voltage, pass short-circuit currents, and withstand high vibration levels to avoid untimely opening or closing, which explains the high mass. Generally speaking, relays used in aeronautics, whether electromechanical or electronic, are bulky and require control electronics to drive them. For electronic relays, the control electronics can be complex because it is necessary to provide protections against current short circuits (therefore providing one or more current sensors) and protections against overvoltages and therefore more sophisticated control electronics than electromechanical relays.This control electronics is also redundant to avoid a simple failure case related to the loss of the control electronics, as this could cause a loss of control of the entire power system. In addition, relays, especially electromechanical relays, have long closing and opening times, which requires them to be controlled in a well-defined sequence and to respect the opening and closing times to avoid short circuits between power sources. While these relays are switching, there is also a risk of power interruption and reset of the equipment.

[0008] The aim of the invention is then to propose an electrical power supply system on board an aircraft, allowing electrical power supply redundancy, while being compatible with the mass constraint linked to a light aircraft.

[0009] To this end, the subject of the invention is an electrical power supply system on board an aircraft, the power supply system being configured to be connected to a rotating machine to receive input electrical energy and to deliver output electrical energy to at least one electrical power supply bus of equipment(s) of the aircraft, the rotating machine being included in a main propulsion system of the aircraft, the power supply system comprising:

[0010] - a primary electric battery, intended to be connected to the rotating machine;

[0011] - at least one secondary electric battery, each configured to power an auxiliary propulsion system of the aircraft;

[0012] - a DC-DC conversion device comprising at least two separate DC-DC conversion modules, each conversion module being connected to a respective electric battery and configured to receive DC energy from said battery and convert it into another DC energy delivered as output;

[0013] - a selection device comprising at least one module for selecting a respective voltage from among several voltages, each selection module being connected to at least two separate conversion modules and configured to select, automatically and according to a selection rule, a voltage from among separate voltages from said conversion modules, then to deliver the selected voltage to a respective power supply bus.

[0014] The electrical power supply system according to the invention then makes it possible to offer electrical power supply redundancy through at least one secondary electric battery in addition to the primary electric battery, also called the main electric battery; through the DC-DC conversion device making it possible to have a predefined voltage value corresponding to each electric battery, despite possible variations in battery voltage;as well as by the selection device making it possible to automatically select one of the separate voltages from the conversion modules of the conversion device, and in doing so, to typically switch from the initially selected DC voltage to another DC voltage in the event of a failure of the voltage source, such as a respective electric battery, associated with the initially selected voltage, for example to switch from the DC voltage corresponding to the primary electric battery to that corresponding to a respective secondary electric battery in the event of a failure of the primary electric battery and / or of the rotating machine to which this primary battery is intended to be connected.;

[0015] Preferably, the or each selection module included in the selection device comprises only passive components, in particular diodes, each selection module typically comprising a diode OR unit (from the English ORing), which makes it possible to further reduce the mass of the electrical power supply system according to the invention.

[0016] According to other advantageous aspects of the invention, the electrical power supply system comprises one or more of the following features, taken individually or in all technically possible combinations:

[0017] - the selection rule is a selection of the highest priority available voltage among the distinct voltages from said conversion modules, following a predefined order of priority of said voltages;

[0018] - the selection rule is a selection of the highest voltage among the distinct voltages from said conversion modules;

[0019] - among the at least two conversion modules connected to each selection module, at least one conversion module is connected to the primary electric battery; and in normal operation of the primary electric battery, the highest voltage among the voltages from said conversion modules is that from the conversion module connected to the primary electric battery; the highest voltage among the voltages from said conversion modules being that from a conversion module connected to a respective secondary electric battery only in the event of a malfunction of the primary electric battery;

[0020] - each selection module is connected to exactly two separate conversion modules, and is configured to select a respective voltage from a first voltage delivered by one of the two conversion modules and a second voltage delivered by the other of the two conversion modules, a value of the first voltage belonging to a first range of values ​​and a value of the second voltage belonging to a second range of values, the second range being separate from the first range. the first range is preferably greater than the second range;

[0021] -each selection module comprises only passive components; each selection module preferably comprising a diode OR unit;

[0022] - the system comprises a single secondary electric battery; the DC-DC conversion device preferably comprising exactly two separate conversion modules;

[0023] - each conversion module comprises two separate output terminals and is configured to deliver a respective voltage at each of its output terminals, and the selection device comprises two selection modules, each being connected to a respective output terminal of each conversion module;

[0024] - the system includes two secondary electric batteries;

[0025] - the DC-DC conversion device comprises four separate conversion modules; two conversion modules are preferably connected to the primary electric battery, and a respective conversion module is connected to each of the two secondary electric batteries;

[0026] - each conversion module has two separate output terminals and is configured to deliver a respective voltage at each of its output terminals, and the selection device has four selection modules, each connected to a respective output terminal of each module of a respective pair of conversion modules; and

[0027] - the DC-DC conversion device comprises three separate conversion modules; a respective conversion module is preferably connected to each of the primary and secondary electric batteries.

[0028] The invention also relates to an aircraft, in particular a drone, comprising:

[0029] - a main propulsion system comprising a rotating machine and a main propeller coupled to the rotating machine;

[0030] - at least one auxiliary propulsion system, each comprising an engine and an auxiliary propeller coupled to the engine;

[0031] - an electrical power supply system as defined above, the power supply system being connected to the rotating machine to receive input electrical energy and configured to deliver output electrical energy to at least one equipment power supply bus(es); the rotating machine preferably being a thermal machine; the motor of the or each auxiliary propulsion system being more preferably an electric motor.

[0032] These characteristics and advantages of the invention will appear more clearly on reading the description which follows, given solely by way of non-limiting example, and made with reference to the appended drawings, in which:

[0033] - Figure 1 is a schematic representation of an aircraft according to the invention, comprising a main propulsion system with a rotating machine, at least one auxiliary propulsion system, and an electrical power supply system connected to the rotating machine to receive an input electrical energy, then to deliver an output electrical energy to one or more equipment of the aircraft;

[0034] - figure 2 is a schematic representation of the electrical power supply system of figure 1, according to a first embodiment;

[0035] - figure 3 is a view similar to that of figure 2, according to a second embodiment;

[0036] - figure 4 is a view similar to that of figure 2, according to a third embodiment; and

[0037] - Figure 5 is a schematic representation of a voltage selection unit.

[0038] In the remainder of the description, the expressions “substantially equal to” and “of the order of” each define a relationship of equality to plus or minus 20%, preferably to plus or minus 10%, and more preferably to plus or minus 5%.

[0039] In Figure 1, an aircraft 5 comprises a main propulsion system 6, at least one auxiliary propulsion system 8, an electrical power supply system 10, and a set 12 of aeronautical equipment. The electrical power supply system 10 is connected, on the one hand, to the main propulsion system 6 to receive an input electrical energy, and on the other hand, to the at least one auxiliary propulsion system 8 and to the set 12 of aeronautical equipment to deliver an output electrical energy to them. In addition, the electrical power supply system 10 is connected to an electronic device 14 external to the aircraft 5, typically via a radio link 16, the electronic device 14 being for example a ground station.

[0040] The aircraft 5 is typically a drone, i.e., an unmanned aerial vehicle (UAV). The drone is, for example, a fixed-wing drone or a rotary-wing drone. Alternatively, the aircraft 5 is a light electric aircraft or a flying electric urban taxi.

[0041] In the example of Figure 1, the aircraft 5 comprises two auxiliary propulsion systems 8, namely a first auxiliary propulsion system 8A and a second auxiliary propulsion system 8B. The first auxiliary propulsion system 8A corresponds for example to a right edge of the aircraft 5, and the second auxiliary propulsion system 8B corresponds for example to a left edge of the aircraft 5.

[0042] The main propulsion system 6 comprises a main propeller 18 and a rotating machine 20 coupled to the main propeller 18 to drive it in rotation. The rotating machine 20 is for example a thermal machine powered by fuel stored in a tank, not shown.

[0043] In addition, the main propulsion system 6 comprises a generator 22 and a control unit 24, the generator 22 being capable of providing electrical energy from the rotation of a shaft, not shown, generated by the rotating machine 20. The generator 22 is for example a permanent magnet generator, also denoted PMG (from the English Permanent Magnet Generate). The control unit 24 is also denoted SGCU (from the English Starter-Generator Control Unit), and is configured to deliver electrical energy, preferably continuous, to the electrical power supply system 10, this from the electrical energy produced by the generator 22. The control unit 24 is also configured to provide electrical energy necessary for starting the rotating machine 20 when the aircraft 5 is on the ground and / or electrical energy necessary for a possible restart of the rotating machine 20 when the aircraft 5 is in flight.

[0044] Each auxiliary propulsion system 8 comprises an auxiliary propeller 26 and a motor 28 coupled to the auxiliary propeller 26 to drive it in rotation. The motor 28 is for example an electric motor.

[0045] In addition, each auxiliary propulsion system 8 comprises a unit 30 for controlling the corresponding motor 28. The control unit 30, visible in FIGS. 2 to 4, is also denoted ESC (Electronic Speed ​​Controller), and is configured to control the speed of the motor 28 from the electrical power supplied by the electrical power system 10.

[0046] Each auxiliary propulsion system 8 is preferably arranged on a respective edge, right or left, of the aircraft 5. Each auxiliary propulsion system 8 is designed to operate at takeoff of the aircraft 5 in order to provide assistance to the propulsion provided by the main propulsion system 6, and also in the event of failure of the main propulsion system 6 to enable safe reaching of a landing site near the aircraft 5. The set 12 of aeronautical equipment, powered by the electrical power supply system 10, comprises a human-machine interface 32, typically including a display screen and one or more aircraft functional control members, not shown; one or more avionics equipment 34, in particular a flight control computer, not shown, also noted FCC (from the English Flight Control Computed); and one or more electrical equipment 36 on board the aircraft.

[0047] The assembly 12 is typically powered by the electrical power supply system 10 via at least one power supply bus 39, and advantageously via several power supply buses 39 arranged in parallel with each other, as shown in FIGS. 2 to 4.

[0048] The electrical power supply system 10 comprises a primary electrical battery 40 connected to the rotating machine 20, for example via the generator 22 and the control unit 24, as shown in FIGS. 1 and 2.

[0049] The generator 22 and the control unit 24 then make it possible to recharge the primary electric battery 40. In particular, the control unit 24 is configured to maintain the primary battery 40 at its maximum state of charge, while respecting a maximum charging current of said battery. When the primary battery 40 is fully charged, the control unit 24 is configured to be able to deliver a maximum current and / or a maximum voltage that does not cause overcharging of the primary battery 40, in order to avoid a risk of thermal runaway of this primary battery 40. The control unit 24 is configured to continue to supply electrical energy to the conversion module(s) 48, described below, which are connected to it.The control unit 24 is also configured to start the rotating machine 20 as described previously, this typically using the primary electric battery 40, the latter then having a minimum charge before takeoff of the aircraft 5.

[0050] The primary electric battery 40 is advantageously a battery having a high power density, such as a power density greater than 3 kW / kg, preferably of the order of 5 kW / kg. The primary electric battery 40 then typically has a low energy density, such as an energy density of between 30 Wh / kg and 100 Wh / kg, and preferably of between 50 Wh / kg and 70 Wh / kg. The primary electric battery 40 is then generally a so-called power type battery, i.e. a battery designed to provide a high quantity of electrical energy over a relatively short period of time. Those skilled in the art will understand that the aforementioned ranges of power density and energy density values ​​correspond to an exemplary embodiment for the primary electric battery 40, and of course that for the same energy density, the higher the power density of the primary electric battery 40, the better.The primary electric battery 40 has, for example, a nominal voltage of the order of 48 V (volts), and the voltage of the primary battery 40 can nevertheless vary between a minimum voltage, for example of the order of 36 V, and a maximum voltage, for example of the order of 54 V.

[0051] The primary electric battery 40 is for example a sodium-ion (Na-ion) battery, or a lithium-ion (Li-ion) battery, or even a lithium-titanate battery.

[0052] The electrical power system 10 comprises at least one secondary electrical battery 42, each configured to power a respective auxiliary propulsion system 8, as shown in FIGS. 2 to 4.

[0053] Each secondary electric battery 42 is advantageously a battery having a high energy density, such as an energy density greater than 150 Wh / kg, and preferably greater than 250 Wh / kg. Each secondary electric battery 42 then typically has a low power density, such as a power density less than 1.5 kW / kg, preferably between 0.5 kW / kg and 1 kW / kg. Each secondary electric battery 42 is then generally a so-called energy-type battery, that is to say a battery designed to store and supply a quantity of electrical energy over an extended period of time and thus offer a high autonomy.Those skilled in the art will understand that the aforementioned ranges of power density and energy density values ​​correspond to an exemplary embodiment for the secondary electric battery(ies) 42, and of course that for the same power density, the higher the energy density of each secondary electric battery 42, the better.

[0054] Each secondary electric battery 42 has, for example, a nominal voltage of the order of 56 V, and the voltage of each secondary battery 42 can nevertheless vary between a minimum voltage, for example of the order of 40 V, and a maximum voltage, for example of the order of 60 V.

[0055] Each secondary electric battery 42 is for example a lithium-ion battery, such as an MC (Manganese Cobalt) lithium-ion battery or an NMC (Nickel Manganese Cobalt) lithium-ion battery.

[0056] Each secondary electric battery 42 is fully charged before takeoff of the aircraft 5, and is then no longer recharged during the flight of the aircraft, each secondary battery 42 being intended to be used during short time periods, for example during takeoff of the aircraft 5 or in emergency mode, as will be explained in more detail later.

[0057] The maximum voltages of the electric batteries 40, 42 are preferably less than 60 V DC for personal safety reasons. The minimum and maximum voltages of the electric batteries 40, 42 are likely to vary from one battery to another.

[0058] In the examples of Figures 2 and 3, the electrical power supply system 10 comprises two secondary electrical batteries 42, namely a first secondary electrical battery 42A and a second secondary electrical battery 42B. The first secondary electrical battery 42A is typically configured to power the first auxiliary propulsion system 8A, corresponding for example to the right edge of the aircraft; and the second secondary electrical battery 42B is typically configured to power the second auxiliary propulsion system 8B, corresponding for example to the left edge of the aircraft.

[0059] In the example of Figure 4, the electrical power system 10 comprises a single secondary electrical battery 42.

[0060] The power supply system 10 comprises a DC-DC conversion device 45, also called a DC-DC (Direct Current) conversion device, comprising at least two separate DC-DC conversion modules 48. Each conversion module 48, also called a DC-DC conversion module, is connected to a respective electric battery 40, 42 and configured to receive DC energy from said battery 40, 42 and convert it into another DC energy delivered as output.

[0061] Each conversion module 48 connected to the primary electric battery 40 is also called a primary conversion module 48P. Each primary conversion module 48P is then configured to convert the voltage of the primary electric battery 40 into a first voltage LU. The first voltage U1 belongs to a first range of values. The first voltage U1 is advantageously of the order of 28 V, and the first range is then preferably substantially centered around 28 V.

[0062] Each conversion module 48 connected to a respective secondary electric battery 42 is also called a secondary conversion module 48S. Each secondary conversion module 48S is then configured to convert the voltage of the respective secondary electric battery 42 into a second voltage U2. The second voltage U2 belongs to a second range of values, advantageously separate from the first range of values ​​associated with the first voltage LU, the second range being preferentially lower than the first range. The second voltage U2 is advantageously of the order of 24 V, and the second range is then preferentially substantially centered around 24 V.

[0063] Each conversion module 48 is known per se, and typically comprises controllable switches, driven to implement said conversion. The controllable switches are for example transistors, such as field effect transistors, in particular MOSFETs (from the English Metal-Oxide-Semiconductor Field-Effect Transistor), optionally each with a diode connected in antiparallel to be bidirectional; or even thyristors.

[0064] The conversion modules 48 are advantageously physically and electrically segregated at the output of each other.

[0065] In addition, each conversion module 48 advantageously includes internal galvanic isolation, not shown, making it possible to prevent propagation of an overvoltage from the input to the output of said conversion module.

[0066] In the example of Figure 2, the conversion device 45 comprises four separate conversion modules 48, namely two primary conversion modules 48P each connected to the primary electric battery 40 and two secondary conversion modules 48S, each secondary conversion module 48S being connected to a respective secondary electric battery 42, i.e. to one of the two secondary batteries 42.

[0067] The electrical power supply system 10 comprises a voltage selection device 50 comprising at least one module 52 for selecting a respective voltage from among several voltages. Each selection module 52 is connected to at least two separate conversion modules 48 and configured to select, automatically and according to a selection rule, a voltage from among the separate voltages from said conversion modules 48, then to deliver the selected voltage to an internal network 38. This internal network 38 is distributed to at least one power supply bus 39, and advantageously via several power supply buses 39 by means of respective protection unit(s) 62, described in more detail below. The internal network 38 is called a high-availability internal network because it is supplied by two input networks whose failures are uncorrelated.

[0068] Each selection module 52 then comprises at least two inputs, each being connected to a respective conversion module 48, and an output connected to the respective power supply bus 39.

[0069] In the examples of Figures 2 to 4, each selection module 52 comprises a selection unit 56 configured to select, automatically and according to the selection rule, the voltage from among the distinct voltages coming from said conversion modules 48, then to deliver the selected voltage to the internal network 38. Each selection unit 56 comprises several inputs and one output, and is then connected by its inputs to said conversion modules 48, and by its output to the internal network 38. In the examples of Figures 2 to 4, each selection module 52 further comprises for each of its inputs, a protection unit, for example in the form of a second protection unit 60 described below; and for each of its outputs connected to the respective power supply buses 39, another protection unit, for example in the form of a third protection unit 62 described below.Each selection unit 56 is then arranged between the second protection units 60 and the third protection unit(s) 62.

[0070] Each selection module 52 is typically connected to both a respective 48P primary conversion module and a respective 48S secondary conversion module, and then configured to select, automatically and according to the selection rule, a voltage from among the first voltage U1 from the 48P primary conversion module and the second voltage U2 from the 48S secondary conversion module.

[0071] In the example of Figure 2, each conversion module 48 has two separate output terminals 54 and is configured to deliver a respective voltage at each of its output terminals 54, and the selection device 50 has four selection modules 52, each selection module 52 being connected to a respective output terminal 54 of each module 48 of a respective pair of conversion modules 48.

[0072] The selection rule is, for example, a selection of the highest voltage among the distinct voltages U1, U2 coming from said conversion modules 48, and each selection module 52 is then configured to automatically select the highest voltage among the distinct voltages coming from said conversion modules 48.

[0073] Each selection unit 56 advantageously comprises only passive components. In the example of FIG. 5, each selection unit 56 then comprises a diode OR unit. The diode OR unit, also known as ORing, comprises two diodes D1, D2, namely a first diode D1 and a second diode D2. The cathodes of the two diodes D1, D2 are typically connected together and connected to a respective internal network 38, the anode of the first diode D1 being connected to the output of a respective primary conversion module 48P to receive the first voltage U1, and the anode of the second diode D2 being connected to the output of a respective secondary conversion module 48S to receive the second voltage U2.

[0074] Alternatively, each selection unit 56 comprises active components, such as transistors and diodes. According to this variant, each selection unit 56 comprises, for example, a logical OR unit, not shown, produced with transistors and diodes. This variant then makes it possible to overcome a residual voltage of the diodes D1, D2, while blocking the reverse current.

[0075] Alternatively, the selection rule is a selection of the highest priority available voltage among the distinct voltages U1, U2 from said conversion modules 48, following a predefined order of priority of said voltages, and each selection module 52 is then configured to automatically select the highest priority available voltage among the distinct voltages from said conversion modules 48.

[0076] According to this variant of the selection rule, each selection module 52, and in particular each selection unit 56, is typically configured to select the highest priority available voltage in the manner described in document FR 2 985 103 A1 published on June 28, 2013.

[0077] According to this variant of the selection rule, typically, the order of priority between the voltages is predefined, and the availability of each of the voltages is defined via a range of respective voltage values ​​for each voltage. Each selection module 52, and in particular each selection unit 56, is then first configured to determine the available voltages among the different distinct voltages from the conversion modules 48 to which it is connected, this by checking for each voltage whether the current value of said voltage belongs to the predefined range of values ​​for said voltage. Each selection module 52, and in particular each selection unit 56, is then configured to select, among the determined available voltages, the one which has the highest priority according to the predefined order of priority for said voltages.

[0078] According to this variant of the selection rule, each conversion module 48 can be freed from the constraint of converting the voltage of the batteries 40 and / or 42 to regulated voltages U1 and U2.

[0079] Furthermore, the galvanic isolation can be ensured by the conversion module 48 and / or by the selection module 52 in the manner described in the document FR 2 985 392 A1 published on July 5, 2013, or in the document EP 2 610 985 A1 published on July 3, 2013.

[0080] Optionally, the power supply system 10 comprises first protection units 58 and / or second protection units 60 and / or third protection units 62 and / or fourth protection units 64 and / or fifth protection units 66.

[0081] Each first protection unit 58 is configured to protect a respective output of a secondary electric battery 42 against an overcurrent. Each first protection unit 58 is also denoted OCP (from the English Over Current Protection).

[0082] Each second protection unit 60 is configured to protect a respective input of a selection module 52 against both an overvoltage and an overcurrent. The second protection units 60 of the same selection module 52 are optionally configured to detect the presence of a reverse current, a sign of failure, and to stop the propagation effects of this failure by opening simultaneously. Each second protection unit 60 is also denoted OVCP (from the English Over Voltage and Current Protection).

[0083] Each third protection unit 62 is configured to protect a respective output of a selection module 52 against an overcurrent. Each third protection unit 62 firstly limits the distributed current, then secondly isolates a power supply bus 39 if the fault persists. Optionally, each third protection unit 62 is configured to limit an overvoltage applied to a power supply bus 39 and limit its propagation to the other power supply buses 39 in parallel with the same selection module 52, this via the use of anti-return diode(s) (assisted or not), or via the use of a voltage limitation system on each of the outputs of the third protection units 62. Each third protection unit 62 is also denoted OCLP (from the English Over Current Limiter and Protection).

[0084] Each fourth protection unit 64 is configured to protect a respective output of a conversion module 48 against an overcurrent. Each fourth protection unit 64 firstly limits the distributed current, then secondly isolates one of the outputs of the respective conversion module 48 if the fault persists. Optionally, each fourth protection unit 64 is configured to limit an overvoltage applied to one of the outputs of the respective conversion module 48 and limit its propagation to the other outputs of said respective conversion module 48, this via the use of anti-return diode(s) (assisted or not), or via the use of a voltage limitation system on each of the outputs of the fourth protection units 64. Each fourth protection unit 64 is also denoted OCLP (from the English Over Current Limiter and Protection).

[0085] The use of an individual conversion module 48 for each voltage U1 or U2 would make it possible to dispense with the use of the fourth protection units 64, the protection then being able to be integrated into the conversion module 48.

[0086] Each fifth protection unit 66 is configured to protect a respective input of a conversion module 48, typically a primary conversion module 48P, against both overvoltage and overcurrent. Each fifth protection unit 66 protects against the occurrence of an overvoltage on the input of the respective primary conversion module 48P, in the event of a regulation failure of the control unit 24 or loss of connection with the primary electric battery 40 (from the English Load-Dump). Each fifth protection unit 66 is optionally configured to detect and protect the respective conversion module 48 from failure, this being in the event of overvoltage or permanent short-circuit / overload on the output of the respective conversion module 48 by stopping a power transfer. Each fifth protection unit 66 is also denoted OVCP (from the English Over Voltage and Current Protection).

[0087] Due to the absence of coupling of the control unit 24 to the secondary electric batteries 42, this fifth protection unit 66 is optional at the input of a secondary conversion module 48S.

[0088] In the examples of Figures 2 to 4, each secondary electric battery 42 has two outputs, one being connected to a respective auxiliary propulsion system and the other being connected to a respective secondary conversion module 48S. According to this optional addition, each secondary electric battery 42 is then equipped with two first protection units 58, namely one for each of its two outputs.

[0089] In the examples of Figures 2 and 4, each conversion module 48 has two outputs, one being connected to a respective selection module 52 and the other being connected to another respective selection module 52. According to this optional addition, each conversion module 48 is then equipped with two fourth protection units 64, namely one for each of its two outputs.

[0090] In the example of Figure 3, the single primary conversion module 48P has four outputs, each connected to a respective selection module 52; and each secondary conversion module 48S has two outputs, one connected to a respective selection module 52 and the other connected to another respective selection module 52. According to this optional addition, the primary conversion module 48P is then equipped with four fourth protection units 64, and each secondary conversion module 48S is equipped with two fourth protection units 64, namely one for each of the respective outputs of said conversion modules 48P, 48S.

[0091] According to this optional addition, each secondary conversion module 48S is then protected at the input by the first protection unit 58 equipping the secondary electric battery 42 to which it is connected, and at the output by the two fourth protection units 64 with which it is equipped.

[0092] According to this optional addition, each selection module 52 is equipped with a second respective protection unit 60 at each of its inputs. Each selection module 52 is then equipped with at least two second protection units 60.

[0093] According to this further optional addition, each selection module 52 is equipped with a respective third protection unit 62 at each of its outputs. Each selection module 52 is then equipped with at least one third protection unit 62, and typically several third protection units 62 in parallel. According to this further optional addition, each primary conversion module 48P is equipped at the input with a respective fifth protection unit 66, in order to protect said module against an overcurrent and / or an overvoltage due to the primary electric battery 40 to which it is connected. Each primary conversion module 48P is then protected at the input by said fifth protection unit 66, and at the output by the fourth protection units 64 with which it is equipped.

[0094] As an optional addition, the power supply system 10 comprises additional converters, not shown, each being configured to be connected to a respective internal network 38 and to then receive the voltage applied to this internal network 38, then convert it into another voltage, typically of a lower value, for example of the order of 5 V or 12 V.

[0095] The person skilled in the art will then observe that in the examples of figures 2 and 3, the aircraft 5 has an architecture distributed along three channels, namely a central channel CC and two lateral channels CA, CB, with a first lateral channel CA corresponding for example to the right edge of the aircraft and a second lateral channel CB corresponding to the left edge of the aircraft. The separations between the channels CA, CB, CC are represented by dotted lines in figures 2 and 3.

[0096] The central channel DC then corresponds to the main propulsion system 6, to the primary electric battery 40 and to each primary conversion module 48P. This central channel DC also corresponds to two central power supply channels VCA, VCB, with a first central channel VCA on the side of the first lateral channel AC (for example on the side of the right edge of the aircraft) and a second central channel VCB on the side of the second lateral channel CB (for example on the side of the left edge of the aircraft). Each central channel VCA, VCB comprises a respective selection module 52 and a respective power supply bus 39.

[0097] Each lateral channel CA, CB then corresponds to a respective auxiliary propulsion system 8, to the secondary electric battery 42 connected to said auxiliary propulsion system 8 and to the secondary conversion module 48S connected to said secondary electric battery 42. Each lateral channel CA, CB also corresponds to a lateral channel VA, VB, the first lateral channel CA corresponding to a first lateral channel VA, for example for the right edge of the aircraft; and the second lateral channel CB corresponding to a second lateral channel VB, for example for the left edge of the aircraft. Each lateral channel VA, VB comprises a respective selection module 52 and a respective power supply bus 39.

[0098] The person skilled in the art will further observe that in the example of Figure 2, the electrical power supply system 10 has an architecture that is symmetrical with respect to a center line, represented in the form of a line of points in Figure 2, with on the one hand the first lateral channel VA and the first central channel VCA (corresponding for example to the right side of the aircraft), and on the other hand the second lateral channel VB and the second central channel VCB (corresponding for example to the left side of the aircraft).

[0099] When the main propulsion system 6 is operating normally, it then recharges the primary electric battery 40, and the primary conversion modules 48P are therefore powered nominally, which then makes it possible to power each power bus 39 with the voltage selected by the respective selection module 52 from the first voltage U1 and the second voltage U2.

[0100] For example, if the selection rule is the selection of the highest voltage, for each selection module 52, when the first voltage U1 is higher than the second voltage U2, i.e. in normal operation of the primary electric battery 40, the first diode D1 is forward biased and allows current to flow from the first voltage U1 to an output voltage delivered to the respective power supply bus 39. The second diode D2 is then reverse biased and blocks current from the second voltage U2 to the output voltage. The output voltage is then substantially equal to the first voltage U1, and the corresponding selection module 52 has selected the first voltage U1 where appropriate.

[0101] In the event of a malfunction of the main propulsion system 6 and / or the primary electric battery 40, the primary conversion modules 48P are no longer supplied nominally, and the secondary conversion modules 48S then make it possible, in emergency mode, to ensure continuity of the electrical power supply to the set 12 of aeronautical equipment of the aircraft via the power supply buses 39.

[0102] Indeed, for each selection module 52, when the second voltage U2 is higher than the first voltage LU, that is to say in the event of a malfunction of the primary electric battery 40, the second diode D2 is forward biased and allows the current to flow from the second voltage U2 to the output voltage. The first diode D1 is then reverse biased and blocks the current from the first voltage U1 to the output voltage. Consequently, the output voltage is then substantially equal to the second voltage U2, and the corresponding selection module 52 has selected the second voltage U2 in this case.

[0103] According to this emergency mode, each secondary electric battery 42 also electrically supplies the auxiliary propulsion system 8 to which it is connected, which then allows the aircraft 5 to land safely.

[0104] As another example, if the selection rule is the selection of the highest priority available voltage, with in addition the highest priority for the first voltage U1, the operation is then similar to that described previously when the selection rule is the selection of the highest voltage.

[0105] The person skilled in the art will observe that if the first voltage U1 and the second voltage U2 are equal, a consumption sharing phenomenon would occur on the two networks causing an unwanted discharge of the respective secondary battery 42. Having the first range and the second range separate makes it possible to ensure the consumption of power on a main network associated with the primary battery 40 as long as it is present without impacting the autonomy available by the respective secondary battery 42.

[0106] The person skilled in the art will also understand that the control unit 24 included in the main propulsion system 6 serves to recharge the primary electric battery 40, and performs a current limitation during the recharging of the primary battery 40, then a voltage regulation when the primary battery 40 is charged. In other words, the control unit 24 is configured to perform a voltage regulation protected by a current limitation.

[0107] Thus, the electrical power supply system 10 according to the invention makes it possible to withstand a double failure, for example a failure of the rotating machine 20 and of the primary battery 40, the electrical power supply of the aircraft 5 then being ensured via the secondary battery(ies) 42, the conversion device 45 and the selection device 50. Indeed, a loss of power supply to the aeronautical equipment of the aircraft 5 and / or all the propulsion systems 6, 8 would be catastrophic events, because the aircraft 5 is intended to fly over populated areas. The electrical power supply system 10 according to the invention then makes it possible to reduce this risk of a catastrophic event.

[0108] Also, in the event of a malfunction of the rotating machine 20, the electrical power supply system 10 according to the invention makes it possible to power the aircraft 5, and in particular each auxiliary propulsion system 8, for a sufficient duration, typically of the order of several tens of minutes, to allow the aircraft 5 to safely reach a nearby landing site.

[0109] The electrical power supply system 10 according to the invention also makes it possible to start the rotating machine 20 when the aircraft 5 is on the ground and / or to restart the rotating machine 20 when the aircraft 5 is in flight, this by supplying - from the primary electric battery 40 to the rotating machine 20 - the electrical energy necessary for its start-up and / or its restart.

[0110] In addition, the electrical power supply system 10 according to the invention also offers protection against voltage and current overloads, in particular via the first, second, third, fourth and fifth protection units 58, 60, 62, 64, 66.

[0111] Figure 3 illustrates a second embodiment of the electrical power supply system 10 according to the invention for which the elements identical to the first embodiment, described previously, are identified by identical references and are not described again.

[0112] According to this second embodiment, the conversion device 45 comprises a single conversion module 48 connected to the primary electric battery 40. In the example of FIG. 3, the conversion device 45 then comprises three conversion modules 48, namely a single primary conversion module 48P and two secondary conversion modules 48S.

[0113] According to this second embodiment, the single primary conversion module 48P is then connected by one of its outputs to each of the selection modules 52. The selection device 50 comprises for example four selection modules 52, in a manner identical to the first embodiment.

[0114] The electrical connections for each of the 48S secondary conversion modules are identical to those described previously for the first embodiment.

[0115] The operation and advantages of the electrical power supply system 10 according to this second embodiment are similar to those described previously for the first embodiment, with the difference that the primary conversion module 48P is not redundant according to this second embodiment, which results in less resistance to failures, but makes it possible to reduce the mass of the electrical power supply system 10, and therefore of the aircraft 5.

[0116] Figure 4 illustrates a third embodiment of the electrical power supply system 10 according to the invention for which the elements identical to the first embodiment, described previously, are identified by identical references and are not described again.

[0117] According to this third embodiment, the aircraft 5 comprises a single auxiliary propulsion system 8, and the electrical power supply system 10 then also comprises a single secondary electric battery 42.

[0118] According to this third embodiment, the conversion device 45 comprises a single conversion module 48 connected to the primary electric battery 40. In the example of FIG. 4, the conversion device 45 then comprises two conversion modules 48, namely a single primary conversion module 48P and a single secondary conversion module 48S.

[0119] According to this third embodiment, the selection device 50 comprises only two selection modules 52, and each conversion module 48 is then connected by one of its outputs to each of the two selection modules 52.

[0120] The operation of the electrical power supply system 10 according to this third embodiment is similar to that described previously for the first embodiment, and is then not described again.

[0121] According to this third embodiment, the electrical power supply system 10 offers less resistance to failures due to less redundancy, by comprising a single primary conversion module 48P, a single secondary battery 42 and a single secondary conversion module 48S, but in return allows a significant reduction in the mass of the electrical power supply system 10, and therefore of the aircraft 5. It is thus understood that the electrical power supply system 10 according to the invention allows electrical power supply redundancy, while being compatible with the mass constraint linked to a light aircraft.

Claims

CLAIMS 1. Electrical power supply system (10) on board an aircraft (5), the power supply system (10) being configured to be connected to a rotating machine (20) to receive input electrical energy and to deliver output electrical energy to at least one electrical power supply bus (39) of equipment(s) of the aircraft (5), the rotating machine (20) being included in a main propulsion system (6) of the aircraft (5), the power supply system (10) comprising: - a primary electric battery (40), intended to be connected to the rotating machine (20); - at least one secondary electric battery (42), each configured to power an auxiliary propulsion system (8) of the aircraft (5); - a DC-DC conversion device (45) comprising at least two separate DC-DC conversion modules (48), each conversion module (48) being connected to a respective electric battery (40, 42) and configured to receive DC energy from said battery (40, 42) and convert it into another DC energy delivered as output; - a selection device (50) comprising at least one module (52) for selecting a respective voltage from among several voltages (U1, U2), each selection module (52) being connected to at least two separate conversion modules (48) and configured to select, automatically and according to a selection rule, a voltage from among separate voltages (U1, U2) from said conversion modules (48), then to deliver the selected voltage to a respective power supply bus (39).

2. System (10) according to claim 1, in which the selection rule is a selection of the highest priority available voltage among the distinct voltages (U1, U2) from said conversion modules (48), following a predefined order of priority of said voltages.

3. System (10) according to claim 1, in which the selection rule is a selection of the highest voltage among the distinct voltages (U1, U2) from said conversion modules (48).

4. System (10) according to claim 3, wherein among the at least two conversion modules (48) connected to each selection module (52), at least one conversion module (48P) is connected to the primary electric battery (40); and in normal operation of the primary electric battery (40), the highest voltage among the voltages (U1, U2) from said conversion modules (48) is that (U1) from the conversion module (48P) connected to the primary electric battery (40); the highest voltage among the voltages (U1, U2) from said conversion modules (48) being that from (U2) a conversion module (48S) connected to a respective secondary electric battery (42) only in the event of a malfunction of the primary electric battery (40).

5. System (10) according to any one of the preceding claims, wherein each selection module (52) is connected to exactly two separate conversion modules (48), and is configured to select a respective voltage from a first voltage (U1) delivered by one (48P) of the two conversion modules (48) and a second voltage (U2) delivered by the other (48S) of the two conversion modules (48), a value of the first voltage (U1) belonging to a first range of values and a value of the second voltage (U2) belonging to a second range of values, the second range being disjoint from the first range. the first range preferably being greater than the second range.

6. System (10) according to any one of the preceding claims, wherein each selection module (52) comprises only passive components; each selection module (52) preferably comprising a diode OR unit.

7. System (10) according to any one of the preceding claims, wherein the system (10) comprises a single secondary electric battery (42); the DC-DC conversion device (45) preferably comprising exactly two separate conversion modules (48).

8. System (10) according to claim 7, in which each conversion module (48) comprises two distinct output terminals (54) and is configured to deliver a respective voltage at each of its output terminals (54), and the selection device (50) comprises two selection modules (52), each being connected to a respective output terminal (54) of each conversion module (48).

9. System (10) according to any one of claims 1 to 6, wherein the system (10) comprises two secondary electric batteries (42).

10. System (10) according to claim 9, wherein the DC-DC conversion device (45) comprises four separate conversion modules (48); two conversion modules (48P) being preferably connected to the primary electric battery (40), and to each of the two secondary electric batteries (42) being connected a respective conversion module (48S).

11. System (10) according to claim 10, wherein each conversion module (48) has two separate output terminals (54) and is configured to deliver a respective voltage at each of its output terminals (54), and the selection device (50) has four selection modules (52), each being connected to a respective output terminal (54) of each module (48) of a respective pair of conversion modules (48P, 48S).

12. System (10) according to claim 9, wherein the DC-DC conversion device (45) comprises three separate conversion modules (48); a respective conversion module (48) preferably being connected to each of the primary (40) and secondary (42) electric batteries.

13. Aircraft (5), in particular drone, comprising: - a main propulsion system (6) comprising a rotating machine (20) and a main propeller (18) coupled to the rotating machine (20); - at least one auxiliary propulsion system (8), each comprising an engine (28) and an auxiliary propeller (26) coupled to the engine (8); - an electrical power supply system (10) according to any one of the preceding claims, the power supply system (10) being connected to the rotating machine (20) to receive input electrical energy and configured to deliver output electrical energy to at least one equipment power supply bus (39).

Images