ELECTRICAL POWER CONVERSION AND TRANSPORT SYSTEM FOR THE INTERNAL HYBRIDIZATION OF AN AIRCRAFT TURBOMACHINE

Abstract

The present invention relates to a system (25) for converting and transporting electrical energy in an aircraft powered by a turbomachine, comprising: - at least one first rotating electrical machine (26) mechanically connected to the low-pressure shaft (12) via a coupling device (27), - a second rotating electrical machine (26') mechanically connected to the high-pressure shaft (11) via an accessory box (17), and - said first electrical power module (28) and said second electrical power module being electrically connected to an internal electrical network (22) of the turbomachine, - a generator (15) coupled with the accessory box (17) and intended to supply electrical power to an aircraft electrical network (16), and - a control device (30) capable of controlling an assistance in starting the turbomachine by at least one rotating electrical machine (26, 26').to compensate for power draws from the generator (15), and to distribute power between power drawn from the high-pressure shaft (11) and power drawn from the low-pressure shaft (12) according to the operating phases of the turbomachine. Figure 3,

Description

Description Title of the invention: CONVERSION AND TRANSPORT OF ELECTRICAL ENERGY FOR INTERNAL HYBRIDIZATION OF A TURBOMACHINE AIRCRAFT

[0001] — The present invention relates to an energy conversion and transport system electric for the internal hybridization of an aircraft turbomachine. The invention finds a particularly advantageous application with propulsion architectures hybrids of civil or military aircraft.

[0002] = As illustrated by [Fig.1], a turbomachine 1 classically comprises a blower 2, one or more stages of compressors, for example a compressor low pressure 3 and a high pressure compressor 4, a combustion chamber 7, a or several stages of turbines, for example a high-pressure turbine 8 and a turbine low pressure 9, and a gas exhaust nozzle (not shown). Typically, The high-pressure turbine 8 drives the high-pressure compressor 4 in rotation by via a first shaft, called the high-pressure shaft 11, while the turbine The low-pressure compressor 9 rotates the low-pressure compressor 3 and the blower 2. via a second shaft, called the low-pressure shaft 12. The low-pressure shaft Pressure 12 is generally housed in the high-pressure shaft 11. A device of FADEC type 14 motor control (for "Full Authority Digital Engine Control") (English) allows monitoring and control of the condition of the turbomachine components 1 and to adjust the various parameters to optimize the performance of the tur- bomachine 1.

[0003] — In order to ensure the supply of power to the aircraft's electrical system in this type of non-hybrid architecture, at least one generator 15 is connected to the upper tree pressure 11 via an accessory box 17 or AGB for "accessory "gearbox" in English, This 17 accessory case usually includes one or several gear trains that are driven in rotation by a sampling mechanical by means of a right-angle drive device on the high-pressure shaft 11. The or the generators 15 can thus electrically supply all the electrical loads connected to the aircraft's electrical network 16, such as the control systems flight systems, de-icing systems, fans, etc…

[0004] A control unit 18 called GCU (for "Generator Control Unit" in English) supervises the operation of the generator 15 in question, notably by ensuring a re- voltage regulation and the establishment of operational diagnostics.

[0005] Such an architecture has the advantage of decoupling the mechanical part propulsive relative to the aircraft electrical network 16. The power available on the aircraft electrical network side 16 is also controlled regardless of the turbomachine 1 operating regime. Within the framework of a hybrid turbomachine 1 architecture, it is possible to implement parallel-connected DC electrical channels as described, for example, in documents WO2020 / 089544 or WO2021 / 099720. An internal hybridization of the turbomachine 1 applying these principles is illustrated in [Fig. 2]. A first DC channel is formed by at least one rotating electrical machine 20.1 mechanically connected to the high-pressure shaft 11. This electrical machine 20.1 is associated with an AC / DC converter 21.1. A second DC channel is formed by at least one rotating electrical machine 20.2 mechanically connected to the low-pressure shaft 12. This electrical machine 20.2 is associated with an AC / DC converter 21.2. The converters 21.1 and 21.2 are electrically connected to an internal electrical network 22 of the turbomachine 1.An electrical converter 24 controlled by a controller 18' provides the interface between the internal electrical network 22 of the turbomachine 1 and the electrical network of the aircraft 16. At least one electrical energy storage subset 23, such as for example a battery or supercapacitors, can optionally be connected to the internal electrical network 22. Such an architecture makes it possible to ensure a starting function of the turbomachine 1 via an electrical power supplied by the electric machines 20.1, 20.2. This architecture also ensures an assistance function of the turbomachine 1 according to which the electrical network injects or draws power on the high pressure shaft 11 or the low pressure shaft 12 according to a reference imposed by the engine control device 14. Due to the electrical coupling between the internal electrical network 22 of the turbomachine 1 and the electrical network of the aircraft 16, a problem of quality of the electrical network 22 on the side of the turbomachine 1, such as the appearance of harmonic pollution, is reflected on the side of the electrical network of the aircraft 16, and vice versa. The invention aims in particular to effectively remedy this drawback by proposing a system for converting and transporting electrical energy in an aircraft propelled by a turbomachine on which power can be drawn or injected via a high-pressure shaft and / or a low-pressure shaft, said system comprises: - at least one first rotating electrical machine mechanically connected to the low-pressure shaft by means of a coupling device, - a first electrical power module associated with said first electrical machine, - a second rotating electrical machine mechanically connected to the upper shaft pressure via an accessory box, and - a second electrical power module associated with said second rotating electrical machine, - said first electrical power module and said second electrical power module being electrically connected to an internal electrical network of the turbomachine, - a generator coupled to the accessory box and intended to electrically power an aircraft electrical network, and - a control device capable of controlling assistance in starting the turbomachine by at least one rotating electrical machine, of compensating for power withdrawals made by the generator, and of carrying out a distribution between power taken from the high-pressure shaft and power taken from the low-pressure shaft according to the operating phases of the turbomachine. The invention thus makes it possible to separate the internal electrical network of the turbomachine from the electrical network of the aircraft, insofar as the generator makes it possible to power the systems connected to the on-board electrical network independently of the engine hybridization system. The invention also makes it possible to maintain independence between the propulsion part and the aircraft by not taking any measurements on the electrical withdrawals on the aircraft part managed by the aircraft manufacturer. The invention also proposes a safe architecture by imposing open-loop operation in the event of a malfunction of the hybridization system. According to one embodiment of the invention, the control device is configured to modify a power distribution coefficient Sp as a function of operating phases of the turbomachine, said power distribution coefficient Sp being defined as follows: Sp = P_hp{P_hp + P_bp) - P_hp being a portion of the high-pressure shaft power passing through the accessory gearbox, - P_bp being a portion of the power of the low pressure shaft passing through the first rotating electrical machine. According to one embodiment of the invention, the portion of the high-pressure shaft power passing through the accessory gearbox is determined from a rotational speed measurement taken by a speed sensor mounted on the high-pressure shaft and a torque value of the high-pressure shaft obtained using a torque estimator. This torque estimator may be positioned between a right-angle drive and an input of the accessory gearbox. The torque estimator may also be located on the generator shaft or on the shaft of the second machine. electrical. It is also possible to use a software torque estimator implementing a dedicated algorithm or mapping. According to one embodiment of the invention, the portion of the low-pressure shaft power passing through the first rotating electrical machine is determined from a current and a voltage of the internal electrical network of the turbomachine. According to one embodiment of the invention, said system comprises two rotating electrical machines mechanically connected to the high-pressure shaft via the accessory housing. According to one embodiment of the invention, said system comprises two rotating electrical machines mechanically connected to the low-pressure shaft. According to one embodiment of the invention, said system comprises two generators mechanically connected to the accessory housing. According to one embodiment of the invention, said system comprises two electrical channels, each formed by two rotating electrical machines and two electrical power modules electrically connected to the internal electrical network of the turbomachine. According to one embodiment of the invention, said system comprises at least one electrical energy storage subset, such as a battery or supercapacitors, electrically connected to the internal electrical network of the turbomachine. According to one embodiment of the invention, said system comprises a generator mechanically connected to the coupling device of the low pressure shaft, said generator being intended to electrically supply the aircraft's electrical network. According to one embodiment of the invention, the generator and the second rotating electrical machine are mounted on the accessory case head-to-tail with respect to each other. The invention also relates to an aircraft comprising an electrical energy conversion and transport system as previously defined. The present invention will be better understood and other features and advantages will become apparent upon reading the following detailed description, which includes embodiments given by way of illustration with reference to the accompanying figures, presented by way of non-limiting examples, which may serve to complete the understanding of the present invention and the explanation of its implementation and, where appropriate, contribute to its definition: [Fig.1] The [Fig.1], already described, is a schematic representation of a non-hybrid turbomachine architecture according to the state of the art equipped with a generator intended to supply an electrical network of the aircraft; [Fig.2] The [Fig.2], already described, is a schematic representation of a hybrid turbomachine architecture equipped with an interface between an internal electrical network of the turbomachine and an electrical network of the aircraft; [Fig.3] The [Fig.3] is a schematic representation of a hybrid turbomachine architecture according to the present invention; [Fig.4] The [Fig.4] is a diagram representing, for different values ​​of a power distribution coefficient, the power levels taken from the high pressure shaft and the low pressure shaft in order to supply electrical loads connected to an aircraft electrical network; [Fig.5][Fig.6][Fig.7][Fig.8][Fig.9] Figures 5 to 9 are schematic representations of variant embodiments of the hybrid turbomachine architecture according to the present invention. It should be noted that the structural and / or functional elements common to the different embodiments have the same references. Thus, unless otherwise stated, such elements have identical structural, dimensional, and material properties. Figure 3 shows a system 25 for converting and transporting electrical energy in an aircraft propelled by a turbomachine 1 already described with reference to Figure 1, on which power can be drawn or injected via the high-pressure shaft 11 and / or the low-pressure shaft 12. The rectangle Env_av encloses the components of the aircraft environment while the rectangle Env_turb encloses the components of the turbomachine environment. The system 25 comprises at least one first rotating electrical machine 26 mechanically connected to the low-pressure shaft 12 via a coupling device 27. The coupling device 27 may, if necessary, incorporate a mechanical function for disengaging the electrical machine 26, particularly in the event of a malfunction of the latter. A first electrical power module 28 is intended to be electrically connected, on the one hand, to the first electrical machine 26 and, on the other hand, to an internal electrical network 22 of the turbomachine 1. The internal electrical network 22 of the turbomachine 1 is preferably a DC electrical network. A second rotating electric machine 26 is mechanically connected to the high-pressure shaft 11 via the accessory box 17. A second electrical power module 28' is electrically connected on one side to the second rotating electric machine 26" and on the other side to the internal electrical network 22 of the turbomachine 1. The first electrical power module 28 and the second electrical power module 28' are thus electrically connected to the internal electrical network 22 of the turbomachine 1. The rotating electrical machines 26, 26 are preferably reversible type electrical machines capable of operating in motor mode and in generator mode. In motor mode, a rotating electrical machine 26, 26" transforms electrical energy taken from the internal electrical network 22 into mechanical energy injected onto the low pressure shaft 12 or the high pressure shaft 11. For this purpose, the electrical power module 28, 28' operates in an inverter mode to transform a direct voltage from the internal electrical network 22 into a polyphase alternating voltage applied to the phases of the corresponding electrical machine 26, 26". In generator mode, a rotating electrical machine 26, 26' transforms mechanical energy taken from the low pressure shaft 12 or the high pressure shaft 11 into electrical energy injected into the internal electrical network 22 of the turbomachine 1. For this purpose, the electrical power module 28, 28" operates in a rectifier mode to transform a polyphase alternating voltage generated by the electrical machine 26, 26' into a direct voltage applied to the internal electrical network 22. An electrical power module 28, 28' thus takes the form of an AC / DC converter. The rotating electrical machines 26, 26' are preferably permanent magnet synchronous machines. Alternatively, the rotating electrical machines 26, 26' could be asynchronous machines or any other type of electrical machine suitable for the application. Furthermore, a generator 15 is coupled with the accessory box 17. The generator 15 is intended to electrically supply an electrical network of the aircraft 16. The generator 15 may take the form of a rotating electrical machine with a wound rotor. The generator 15 and the second rotating electric machine 26 are mounted on the accessory box 17 in a back-to-back configuration. This arrangement maximizes stiffness and minimizes response times. It also allows the overall dimensions to be adapted to the application. Alternatively, the generator 15 and the second rotating electric machine 26 can be mounted on the same side of the accessory box 17. A control device 30 is capable of operating a starting assistance system for the turbomachine 1 by at least one rotating electrical machine 26, 26', of compensating for power draws from the generator 15, and of distributing power between power drawn from the high-pressure shaft 11 and power drawn from the low-pressure shaft 12 according to the operating phases of the turbomachine 1 Advantageously, the control device 30 is configured to modify a power distribution coefficient Sp as a function of operating phases of the turbomachine 1 defined as follows: Sp = P_hp{P_hp + P_bp) - P_hp being a portion of the power of the high-pressure shaft 11 passing through the accessory gearbox 17, - P_bp being a portion of the power P_bp of the low pressure shaft 12 passing through the first rotating electrical machine 26. The portion of the power P_hp of the high-pressure shaft 11 passing through the accessory gearbox 17 is determined from a rotational speed measurement N2 taken by a speed sensor 32 mounted on the high-pressure shaft 11 and a torque value C_hp of the high-pressure shaft 11 obtained using a torque estimator. The torque estimator may be positioned between a right-angle drive and an input of the accessory gearbox 17. The torque estimator may also be located on the generator shaft 15 or on the shaft of the second electric machine 26. Alternatively, a software torque estimator implementing a dedicated algorithm or mapping may be used. The portion of the power P_bp of the low pressure shaft 12 passing through the rotating electrical machine 26 is determined from a current I and a voltage U of the internal electrical network 22. Figure 4 illustrates different values ​​of the power distribution coefficient Sp during different operating phases of the turbomachine 1. During an operating phase Ph_1, the value of the power distribution coefficient Sp is 1 so that all the power P_ch consumed by the electrical loads connected to the aircraft electrical network 16 is taken from the high-pressure shaft 11 via the generator 15. We then find ourselves in an operating configuration equivalent to that of a non-hybrid architecture. During an operating phase Ph_2, the value of the power distribution coefficient Sp is 0.5, so that the power P_ch consumed by the electrical loads connected to the aircraft electrical network 16 is distributed equally between a power P_hp taken from the high pressure shaft 11 and a power P_bp taken from the low pressure shaft 12. During an operating phase Ph_3, the value of the power distribution coefficient Sp is O so that all the power P_ch consumed by the electrical loads connected to the aircraft electrical network 16 is taken from the low pressure shaft 12. The power P_bp drawn from the low-pressure shaft 12 is converted into electrical power P_inj on the internal electrical network 22 of the turbomachine 1 by the electric machine 26 and the associated module 28. This electrical power is then converted by the electric machine 26' and the associated module 28 into mechanical power applied to the accessory box 17 to which the generator 15 is connected. Of course, these three scenarios have been presented to facilitate understanding of the invention. The power distribution coefficient Sp may take intermediate values ​​from those shown in the diagram. To perform torque control C_cons of the electric machine 28', a power calculation module 33 receives as input a power distribution coefficient value Sp, a power value injected into the internal electrical network 22 of the turbomachine 1, and a power value P_hp of the high-pressure shaft 11. Once the torque setpoint C_cons is determined from these values, the power control module 28' can drive the rotating electric machine 26' in PWM (Pulse Width Modulation) mode or in full-wave mode. The chosen control mode for the electric machines 26 and 26' depends on the application. In the event of a malfunction of the hybridization system, an open-loop control with imposed sampling on the low-pressure shaft 12 or with imposed sampling on the high-pressure shaft 11 is preferably implemented. In the case where the rotating electrical machines 26, 26" are used solely to draw power from the low-pressure shaft 12 in order to inject power into the aircraft's electrical network 16, it is possible to simplify the hybrid architecture configuration. Indeed, in this case, the electrical machine 26 can function as a generator, while the associated power module 28 can simply act as a voltage rectifier bridge, a more economical function than a power module that also functions as an inverter. In the embodiment of [Fig. 5], the system 25 comprises two reversible rotating electrical machines 26' mechanically connected to the high-pressure shaft 11 via the accessory housing 17. Each rotating electrical machine 26" is associated with a power electrical module 28' connected to the internal electrical network 22 of the turbomachine 1. Alternatively, the system 25 comprises two rotating electrical machines 26 mechanically connected to the low-pressure shaft 12. In the embodiment of [Fig.6], the system 25 includes two generators 15 mechanically connected to the accessory box 17 and intended to supply the electrical network of the aircraft 16. In the embodiment of [Fig.7], the system 25 comprises two electrical channels 34.1, 34.2 each formed by two rotating electrical machines 26, 26" and two electrical power modules 28, 28' electrically connected to the internal electrical network 22 of the turbomachine 1. In addition, two generators 15 intended to supply the on-board network are mechanically connected to the accessory box 17. In the embodiment of [Fig.8], system 25 includes at least one sub assembly 35 of electrical energy storage, such as a battery or supercapacitors, electrically connected to the internal electrical network 22 of the turbomachine 1. In the embodiment of [Fig. 9], the system 25 includes a generator 15" mechanically connected to the coupling device 27 of the low-pressure shaft 12. The generator 15' is associated with a measuring sensor 32' mounted on the low-pressure shaft 12 to determine, among other things, its rotational speed. The generator 15' is intended to supply electrical power to the aircraft's electrical network 16. Of course, the different features, variants and / or embodiments of the present invention can be combined with each other in various ways as long as they are not incompatible or mutually exclusive. Furthermore, the invention is not limited to the embodiments described above and provided solely by way of example. It encompasses various modifications, alternative forms, and other variations that a person skilled in the art may envision within the scope of the present invention, and in particular all combinations of the different modes of operation described above, which may be considered separately or in combination.

Claims

Demands

1. System (25) for converting ct to dc for transporting electrical energy in a aircraft powered by a turbomachine (1) from which one can extract or inject power via a high-pressure shaft (11) and / or a shaft low pressure (12), characterized in that said system (25) comprises: - at least one first rotating electrical machine (26) mechanically connected nimically to the low-pressure shaft (12) via a coupling device (27), - a first electrical power module (28) associated with said first electric machine (26), - a second rotating electrical machine (26") mechanically connected connected directly to the high-pressure shaft (11) via a housing accessories (17), and - a second electrical power module (28') associated with said second rotating electric machine (26"), - said first electrical power module (28) and said second electrical power module (28") being electrically connected to an internal electrical network (22) of the turbomachine (1), - a generator (15) coupled with the accessory box (17) and intended to supply electrical power to an aircraft electrical network (16), And - a control device (30) capable of operating an assistance to starting the turbomachine (1) by at least one electric machine rotating (26, 26"), to compensate for power withdrawals performed by the generator (15), and to perform a distribution between a power taken from the high-pressure shaft (11) and a power taken from the low-pressure shaft (12) according to operating phases operation of the turbomachine (1).

2. System according to claim 1, characterized in that the device of control (30) is configured to modify a distribution coefficient of power Sp as a function of the operating phases of the turbine bomachine (1), said power distribution coefficient Sp being defined as follows: Sp = P_hp / (P_hp + P_bp) - P_hp being a portion of the power of the high-pressure shaft (11) passing through the accessory box (17), - P_bp being a portion of the power of the low-pressure shaft (12) passing through the first rotating electrical machine (26).

3. System according to claim 2, characterized in that the portion of the power {P_hp) of the high-pressure shaft (11) passing through the housing accessories (17) is determined from a speed measurement of rotation (N2) performed by a speed sensor (32) mounted on the shaft high pressure (11) and a high pressure shaft torque value (11) obtained using a torque estimator.

4. System according to claim 2 or 3, characterized in that the portion of the power (P_bp) of the low-pressure shaft (12) passing through the The first rotating electrical machine (26) is determined from a current (I) and a voltage (U) of the internal electrical network (22) of the turbomachine (1).

5. A system according to any one of claims 1 to 4, characterized in which comprises two rotating electrical machines (26) connected me- mechanically to the high-pressure shaft (11) via the housing accessories (17).

6. System according to any one of claims | to 5, characterized in which comprises two rotating electrical machines (26) connected me- mechanically to the low pressure shaft (12).

7. A system according to any one of claims 1 to 6, characterized in which comprises two generators (15) mechanically connected to the accessory case (17).

8. A system according to any one of claims 1 to 7, characterized in which comprises two electrical channels (34.1, 34.2) each formed by two rotating electrical machines (26, 26") and two modules electrical power (28, 28") electrically connected to the grid internal electrical (22) of the turbomachine (1).

9. A system according to any one of claims 1 to 8, characterized in which includes at least one subset (35) of energy storage electrical, such as a battery or supercapacitors, electrically connected trically on the internal electrical network (22) of the turbomachine (1).

10. | System according to any one of claims 1 to 9, characterized in which includes a generator (15") mechanically connected to the coupling device (27) of the low-pressure shaft (12), said ge- neratrice (15') being intended to electrically supply the network electrical (16) of the aircraft.

11. | System according to any one of claims 1 to 10, characterized in what the generator (15) and the second rotating electrical machine (26") are mounted on the accessory housing (17) head-to-tail, one by relationship to the other.

12. Aircraft incorporating a power conversion and transport system electrical as defined according to any one of the preceding claims previous.