Hybrid electrical power supply system for electrical equipment of an aircraft from a turbomachine

Abstract

Hybrid electrical power supply system for electrical equipment of an aircraft from a turbomachine of the aircraft comprising a high-pressure shaft 5 and a low pressure shaft 3, the power supply system comprising a first electrical machine 9 intended to be mechanically connected to the high-pressure shaft of the turbomachine and comprising two first independent stator windings, and - a second electrical machine 10 intended to be mechanically connected to the low pressure shaft of the turbomachine and comprising two second independent stator windings. The second electrical machine is an electrical wound-rotor synchronous machine comprising an electrical outlet intended to be connected directly to a main alternating current network 11 of the aircraft.

Description

Title of the invention: Hybrid electrical power supply system for electrical equipment of an aircraft from a turbomachine Technical Field The invention relates to the internal hybridisation of a turbomachine for electrified aircraft, and more particularly to a hybrid electrical power supply system for electrical equipment of an aircraft from a turbomachine. Prior Art Climate change is a major focus for many legislative and regulatory bodies worldwide. In fact, various restrictions on carbon emissions have been, are or will be adopted by various states. In particular, an ambitious standard applies both to new types of planes and also to those already in circulation, requiring implementation of technological solutions to bring them into line with current regulations. For several years now, civil aviation has been working to contribute towards combating climate change. Technological research efforts have already led to significant improvements in the environmental performance of aircraft. The Applicant is considering the impacting factors in all phases of design and development to obtain aeronautical components and products that consume less energy, are environmentally friendlier and the integration and use of which in civil aviation have moderate environmental consequences with the aim of improving the energy efficiency of aircraft. As a consequence, the Applicant is constantly working to reduce its negative impact on the climate by exploiting green development and manufacturing processes to minimise greenhouse gases emissions as far as possible to reduce the environmental footprint of its business. This sustained research and development work is focused on new generations of aircraft engines, making aircraft lighter, particularly through the materials used and lighter on-board equipment, developing the use of electrical technologies for propulsion and, as an essential complement to technological progress, aeronautical biofuels. In terms of aircraft propulsion, the hybridisation of a turbomachine is achieved by means of an electric system forming the interface between the mechanical shafts of the turbomachine and the aircraft electrical network. This electrical system must ensure the following functions, specifically power generation, turbomachine assistance, start-up of the turbomachine, and reconfiguration as a function of faults. Such an electrical system must provide an electrical network by ensuring network quality and stability to the electrical distribution of the aircraft by drawing power on the low-pressure and high-pressure components of the turbomachine (power generation). It must also inject or draw power on the high-pressure and / or low-pressure components as a function of the instructions received by the turbomachine computer (assistance of the turbomachine). During start-up of the turbomachine it must also provide sufficient mechanical power on the high-pressure shaft, this power capable of originating from a source external to the network (turbomachine start-up). Finally, the system must allow reconfiguration as a function of internal or external fault. The distribution architecture within the scope of the invention corresponds to an internal hybridisation turbomachine scheme based on continuous power channels (DC channels) connected in parallel. These DC channels comprise an electrical machine (often a synchronous permanent magnet machine) linked to a controlled power electronics converter (with a channel for the high-pressure shaft and a channel for the low-pressure shaft) powering a DC bar. There are power converters interfacing between the electrical machines and the network or DC bar. An external source such as an auxiliary power unit (APU) or a battery can also be connected to this DC bar. Expressed in more detail, such an internal hybrid system with DC power bus conventionally comprises a first machine, of synchronous permanent magnet type (PMG), composed of two independent stator windings on the high-pressure shaft of the turbomachine, a first DC / AC converter connected to each stator winding of the first machine, a second machine of synchronous permanent magnet type (PMG) comprising 2 independent stator windings on the low-pressure shaft of the turbomachine, a second bidirectional DC / AC converter connected to each stator winding of the second machine, and two electrical distribution boxes (PDMU) placing the first converter and the second converter in parallel on the same DC power bus and distributing the high-voltage direct electrical energy (HVDC) to the aircraft and to the loads internal to the turbomachine. The term bidirectional DC / AC converter by stator winding covers both a bidirectional converter box and two unidirectional converter boxes (in two opposite directions). This type of internal hybridisation architecture for the turbomachine, placing the electrical machines and their power electronics converters for the high- and low-pressure components used as DC sources in parallel on the same HVDC network, is a solution that represents a technological breakthrough for aircraft manufacturers compared to current aircraft electrical networks. Since the electrical network architectures of future aircraft are unknown, the question of interface between the turbomachine hybridisation network and the aircraft electrical network is crucial. Disclosure of the Invention The aim of the invention, which is particularly advantageous with a view to reducing the environmental impact of aircraft, is to provide a hybrid electrical system for supplying electrical equipment on an aircraft from a turbomachine, enabling hybridisation functions and main electrical generation of the aircraft to be carried out with a known electrical network already in service with aircraft manufacturers, to limit risks and find a realistic interface for aircraft manufacturers between the hybrid electrical network and the aircraft electrical network. In a first aim of the invention, a hybrid electrical power supply system for electrical equipment of an aircraft from a turbomachine of the aircraft is proposed comprising a high-pressure shaft and a low-pressure shaft, the power supply system comprising : - a first electrical machine intended to be mechanically connected to the high-pressure shaft of the turbomachine and comprising two first independent stator windings, and - a second electrical machine intended to be mechanically connected to the low-pressure shaft of the turbomachine and comprising two second independent stator windings. According to a general characteristic of the invention, the second electrical machine is an electrical wound-rotor synchronous machine comprising an electrical outlet intended to be connected directly to a main alternating current network of the aircraft. This architecture makes it possible to : - reduce technological leaps by coupling to the low-pressure shaft an electrical machine of electrical rotor-wound synchronous machine type, such as a variablefrequency electrical generator, which has the secondary advantage of being known and currently in use on aircraft in service, - provide an alternative electrical network (AC) known and in service on aircraft for reusing known electrical loads, - starting up the turbomachine from the aircraft network by injecting power on the high-pressure shaft via the network AC of the aircraft, and - performing hybridisation functions of the turbomachine by powering turbomachine loads via the AC network of the aircraft and by having the capacity to draw power on the high-pressure shaft or on the low-pressure shaft. In a first aspect of the system, the first electrical machine can be an electrical synchronous machine with permanent magnets, and for each first stator winding the system can also comprise an electrical distribution control unit and a bidirectional DC-AC electrical power converter electrically connected between the first stator winding and the electrical distribution control unit, said electrical distribution unit being connected at output to electrical loads of the aircraft. This architecture therefore provides the aircraft with two different types of electrical networks. In addition to the known alternating current (AC) network in service on aircraft, enabling known loads to be reused, it also provides a high-voltage direct current (HVDC) network distributed by the electrical distribution control unit, or PDMU for power distribution monitoring unit. This architecture also performs hybridisation functions of the turbomachine by powering the loads of the turbomachine via the PDMU and by being capable of drawing power on the high-pressure shaft, but in this configuration also injecting power onto the high-pressure shaft. In a second aspect of the system, the second electrical machine can be lacking electrical connection to a power distribution control unit. In this embodiment, the second electrical machine intended to be coupled to the low-pressure shaft, and which comprises an electrical outlet intended to be connected directly to a main alternating current network of the aircraft, is accordingly solely dedicated to the main electrical AC generation of the aircraft. An HVDC network can be reconstituted on the aircraft side and interface with the local HVDC network of the turbomachine via the PDMU. In a third aspect of the system, for each second stator winding the system can also comprise a unidirectional AC / DC electrical power converter connected electrically between the second electrical machine and the corresponding power distribution control unit. In this configuration system, the power can be drawn solely on the low-pressure shaft. Such drawing allows powering the aircraft AC network and powering the HVDC aircraft network and or the loads of the turbomachine via the unidirectional AC / DC electrical power converter and the PDMU. In a fourth aspect of the system, each power distribution control unit can be intended to be connected also to a main direct current power network of the aircraft. In this way, the architecture powers local domestic loads to the turbomachine via the PDMU and provides the option of drawing or injecting power to the high-pressure shaft via the local HVDC network. In a fifth aspect of the system, each power distribution control unit can be lacking an electrical connection to a main direct current power network of the aircraft. This configuration therefore allows powering the domestic loads of the turbomachine via the local HVDC network, and draws or injects power on the high-pressure shaft via the local HVDC network, and stars up the turbomachine by way of injecting power onto the high-pressure shaft via the alternating electrical network originating from the aircraft. In a sixth aspect of the system, the first electrical machine can be an electrical rotorwound synchronous machine, such as a variable-frequency generator (VFG) for example comprising 2 independent stator windings distributed directly to the aircraft network. This configuration eliminates the need for power converters and the HVDC network. According to another aspect of the invention, a turbomachine of an aircraft is proposed, comprising a high-pressure shaft and a low-pressure shaft and a power supply system for electrical equipment of an aircraft such as defined previously. According to yet another aspect of the invention an aircraft is proposed, comprising at least one turbomachine such as defined previously. Brief Description of the Drawings The invention will be better understood from the following description given by way of indication but non-limiting in reference to the attached drawings in which: [Figure 1] Figure 1 shows a turbomachine of an aircraft fitted with a power supply system for electrical equipment of the aircraft according to a first embodiment of the invention. [Figure 2] Figure 2 illustrates a turbomachine of an aircraft fitted with a power supply system for electrical equipment of the aircraft according to a second embodiment of the invention. [Figure 3] Figure 3 illustrates a turbomachine of an aircraft fitted with a power supply system for electrical equipment of the aircraft according to a third embodiment of the invention. [Figure 4] Figure 4 illustrates a turbomachine of an aircraft fitted with a power supply system for electrical equipment of the aircraft according to a fourth embodiment of the invention. Description of Embodiments Figure 1 illustrates a turbomachine of an aircraft fitted with a power supply system for electrical equipment of the aircraft according to a first embodiment of the invention. The turbomachine 1 comprises a fan 2 coupled to a low-pressure shaft 3, the fan 2 acting as a low-pressure compressor and the low-pressure shaft 3 also being coupled to a low-pressure turbine 4. The turbomachine 1 also comprises a high- pressure compressor 5 and a high-pressure turbine 6 coupled to a high-pressure shaft 7. The turbomachine 1 also comprises an electrical power supply system 8 which is a hybrid power supply system for electrical equipment. The power supply system 8 comprises a first electrical machine 9 and a second electrical machine 10. The first electrical machine 9 is mechanically connected to the high-pressure shaft 7 of the turbomachine 1 and comprises two first independent stator windings. The first electrical machine 9 is a synchronous electrical machine with permanent magnets. The second electrical machine 10 is mechanically connected to the low-pressure shaft 3 of the turbomachine 1 and comprises two second independent stator windings. The second electrical machine 10 is a synchronous rotor-wound electrical machine comprising an electrical outlet connected directly to a main alternating current network 11 of the aircraft. For each first stator winding of the first electrical machine 9, the power supply system 8 comprises a power distribution control unit 12 (PDMU) and a bi-directional DC-AC electrical power converter 13 connected electrically between the first stator winding and the power distribution control unit 12, the electrical distribution unit12 being connected at output to a high-voltage DC power network 16 of the aircraft and to electrical loads 14 of the turbomachine 1. For each second stator winding of the second electrical machine 10, the power supply system 8 comprises a unidirectional AC / DC electrical power converter 15 connected electrically between the second electrical machine 10 and the corresponding power distribution control unit 12. Each power distribution control unit 12 is also connected to a main direct current power network 16 (HVDC) of the aircraft. In this way, the architecture powers local domestic Ioads14 of the turbomachine 1 via the PDMU 12 and provides the possibility of drawing from or injecting power onto the high-pressure shaft 7 via the local HVDC network. Figure 2 illustrates a turbomachine of an aircraft fitted with a power supply system for electrical equipment of the aircraft according to a second embodiment of the invention. The second embodiment illustrated in figure 2 differs from the first embodiment illustrated in figure 1 in that each power distribution control unit 12 is lacking an electrical connection to a main direct current power network (HVDC) of the aircraft. This configuration enables domestic loads of the turbomachine to be powered via the local HVDC network, and power to be drawn from or injected onto the high-pressure shaft via the local HVDC network, and the turbomachine to be started by injecting power onto the high-pressure shaft via the alternating current network originating from the aircraft. Figure 3 illustrates a turbomachine of an aircraft fitted with a power supply system for electrical equipment of the aircraft according to a third embodiment of the invention. The third embodiment differs from the first embodiment illustrated in figure 1 in that the second electrical machine 10 is lacking electrical connection to the power distribution control unit 12. In this third embodiment, the second electrical machine 10 coupled to the low-pressure shaft 3 is dedicated solely to the main AC power generation of the aircraft. In an alternative, an HVDC network can be reconstituted on the aircraft side and be interfaced with the local HVDC network of the turbomachine via the PDMU. Figure 4 illustrates a turbomachine of an aircraft fitted with a power supply system for electrical equipment of the aircraft according to a fourth embodiment of the invention. The fourth embodiment differs from the first embodiment illustrated in figure 1 in that the first electrical machine 90 is a rotor-wound electrical synchronous machine 90, such as a variable-frequency generator (VFG) for example comprising 2 independent stator windings directly distributed to the aircraft network, and in that the electrical power supply system 8 is lacking a power distribution control unit, a bidirectional AC / DC electrical power converter and a unidirectional AC / DC electrical power converter. This configuration eliminates the need for power converters and the HVDC network. The invention therefore provides a hybrid electrical power supply system for electrical equipment of an aircraft from a turbomachine by performing the hybridisation functions and the main generation of the aircraft with a known electrical network already in service with aircraft manufacturers to limit risks and find a realistic interface for aircraft manufacturers between the hybrid electrical network and the aircraft electrical network.

Claims

1. A hybrid electrical power supply system (8) for electrical equipment of an aircraft from a turbomachine (1) of the aircraft comprising a high-pressure shaft (7) and a low-pressure shaft (3), the power supply system (8) comprising :- a first electrical machine (9, 90) intended to be mechanically connected to the high-pressure shaft (7) of the turbomachine (1) and comprising two first independent stator windings, and- a second electrical machine (10) intended to be mechanically connected to the low-pressure shaft (3) of the turbomachine (1) and comprising two second independent stator windings, characterised in that:- the second electrical machine (10) is a synchronous rotor-wound electrical machine comprising an electrical outlet intended to be connected directly to a main alternating current network (11) of the aircraft,- the first electrical machine (9) is a synchronous electrical machine with permanent magnets, andthe system (8) also comprises :- for each first stator winding, a power distribution control unit (12) and an electrical converter of bidirectional continuous-alternating power (13) connected electrically between the first stator winding and the power distribution control unit (12), said electrical distribution unit(12) being connected at output to a high-voltage DC network (HVDC) of the aircraft (16) and to electrical loads (14) of the turbomachine (1), and- for each second stator winding, a unidirectional AC / DC electrical power converter (15) connected electrically between the second electrical machine (10) and the corresponding power distribution control unit (12).

2. The system (8) according to claim 1, wherein the second electrical machine (10) is lacking electrical connection to a power distribution control unit (12).

3. The system (8) according to any one of claims 1 or 2, wherein each power distribution control unit (12) is intended to be also connected to a main direct current power network (16) of the aircraft.

4. The system (8) according to any one of claims 1 or 2, wherein each5 power distribution control unit (12) is lacking an electrical connection to a main direct current power network of the aircraft.

5. The turbomachine (1) of an aircraft comprising a high-pressure shaft (7) and a low-pressure shaft (3) and a power supply system for electrical equipment (8) of an aircraft according to any one of claims 1 to 4.io

6. An aircraft comprising at least one turbomachine (1) according to claim 5.12

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