Hybrid electric system for powering electrical equipment of an aircraft from a turbomachine

Abstract

Hybrid electric system for powering electrical equipment of an aircraft from a turbomachine Hybrid electric system (8) for powering 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 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. The second electrical machine (10) is a wound rotor synchronous electrical machine comprising an electrical output intended to be directly connected to a main AC electrical network (11) of the aircraft. Figure for the abstract: Fig. 1

Description

Title of the invention: Hybrid electric system for powering electrical equipment of an aircraft from a turbomachine Technical field

[0001] The invention relates to the internal hybridization of a turbomachine for electrified aircraft, and more particularly to a hybrid electric system for powering electrical equipment of an aircraft from a turbomachine. Prior art

[0002] Climate change is a major concern for many legislative and regulatory bodies around the world. Indeed, various restrictions on carbon emissions have been, are being or will be adopted by various states. In particular, an ambitious standard applies both to new types of aircraft and those in circulation requiring the implementation of technological solutions in order to make them compliant with current regulations. Civil aviation has been mobilizing for several years now to make a contribution to the fight against climate change.

[0003] Technological research efforts have already made it possible to significantly improve the environmental performance of aircraft. The Applicant takes into consideration the impact factors in all phases of design and development to obtain less energy-intensive, more environmentally friendly aeronautical components and products whose integration and use in civil aviation have moderate environmental consequences with the aim of improving the energy efficiency of aircraft.

[0004] Consequently, the Applicant is constantly working to reduce its negative climate impact by using methods and operating virtuous development and manufacturing processes and minimizing greenhouse gas emissions to the minimum possible in order to reduce the environmental footprint of its activity.

[0005] This sustained research and development work covers new generations of aircraft engines, the lightening of aircraft, particularly through the materials used and lighter on-board equipment, the development of the use of electrical technologies to ensure propulsion, and, as essential complements to technological progress, aeronautical biofuels.

[0006] In the context of aircraft propulsion, the hybridization of a turbomachine is carried out using an electrical system interfacing between the mechanical shafts of the turbomachine and the aircraft's electrical network. This electrical system must be able to ensure the following functions, namely, power generation, turbomachine assistance, turbomachine start-up, and reconfiguration in the event of failures.

[0007] Such an electrical system must make it possible to provide an electrical network while ensuring network quality and stability towards the electrical distribution of the aircraft by taking power from the low pressure and high pressure parts of the turbomachine (power generation). It must also make it possible to inject or take power from the high pressure and / or low pressure parts according to the instructions received by the turbomachine computer (turbomachine assistance). It must also make it possible, when starting the turbomachine, to provide sufficient mechanical power to the high pressure shaft, this power being able to come from a source external to the network (turbomachine start). Finally, the system must allow reconfiguration according to an internal or external failure.

[0008] The distribution architecture in the context of the invention corresponds to an internal turbomachine hybridization scheme based on continuous power channels (DC channels) connected in parallel. These DC channels comprise an electrical machine (often a permanent magnet synchronous machine) associated with a controlled power electronics converter (with a channel for the high pressure shaft and a channel for the low pressure shaft) supplying a DC bar. There are thus power converters at the interface 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.

[0009] In more detail, such an internal hybridization system with a DC power bus conventionally comprises a first machine, of the permanent magnet synchronous (PMG) type, 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 the permanent magnet synchronous (PMG) type, composed of 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 units (PDMU) placing the first converter and the second converter in parallel on the same DC power bus and making it possible to distribute the high voltage direct current (HVDC) electrical energy to the aircraft and to the loads internal to the turbomachine.The expression bidirectional DC / AC converter by stator winding covers both a bidirectional converter box or two unidirectional converter boxes (in two opposite directions).

[0010] Such an internal hybridization architecture of the turbomachine, placing in parallel on the same HVDC network, the electrical machines and their power electronic converters of the high and low pressure parts used as DC sources, is a solution which represents a technological breakthrough for aircraft manufacturers compared to current aircraft electrical networks. Not knowing the electrical network architectures of future aircraft, the question of the interface between the hybridization network of the turbomachine and the aircraft electrical network is essential. Statement of the invention

[0011] The invention, which is particularly advantageous for the purpose of reducing the environmental impact of aircraft, aims to provide a hybrid electric system for supplying electrical equipment of an aircraft from a turbomachine making it possible to carry out the hybridization functions and the main electrical generation of the aircraft with a known electrical network already in service with aircraft manufacturers in order to limit the risks and find a realistic interface for aircraft manufacturers between the hybrid electric network and the aircraft electrical network.

[0012] In a first object of the invention, there is proposed a hybrid electric system for supplying electrical equipment of an aircraft from a turbomachine of the aircraft comprising a high pressure shaft and a low pressure shaft, the 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.

[0013] According to a general characteristic of the invention, the second electrical machine is a wound rotor synchronous electrical machine comprising an electrical output intended to be directly connected to a main alternating electrical network of the aircraft.

[0014] This architecture allows both: - to reduce technological leaps by coupling to the low pressure shaft an electrical machine of the wound rotor synchronous electrical machine type, such as a variable frequency electric generator, which has the secondary advantage of being known and currently used on aircraft in service,

[0015] - to provide a known and in-service alternating current (AC) electrical network on aircraft allowing the reuse of known electrical charges,

[0016] - to start the turbomachine from the aircraft network by injecting power into the high pressure shaft through the aircraft's AC network, and

[0017] - to carry out the hybridization functions of the Turbomachine by supplying the Turbomachine loads via the aircraft's AC network and having the ability to draw power from the high pressure shaft or the low pressure shaft.

[0018] In a first aspect of the system, the first electrical machine may be a permanent magnet synchronous electrical machine, and the system may further comprise, for each first stator winding, 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.

[0019] This architecture thus makes it possible to provide two different types of electrical networks to the aircraft. In addition to the known alternating current (AC) network in service on aircraft allowing known loads to be reused, it also provides a high voltage direct current (HVDC) network distributed by the power distribution control unit, or PDMU.

[0020] This architecture also makes it possible to perform the hybridization functions of the turbomachine by supplying the loads of the turbomachine via the PDMU and having the capacity to take power from the high pressure shaft, but in addition, in this configuration, to inject power into the high pressure shaft.

[0021] In a second aspect of the system, the second electrical machine may be devoid of electrical connection to an electrical distribution control unit.

[0022] In this embodiment, the second electrical machine intended to be coupled to the low pressure shaft and which comprises an electrical output intended to be directly connected to a main AC electrical network of the aircraft, is thus solely dedicated to the main electrical generation of the aircraft in AC.

[0023] An HVDC network can be reconstituted on the aircraft side and interface with the local HVDC network of the turbomachine via the PDMU.

[0024] In a third aspect of the system, the system may further comprise, for each second stator winding, a unidirectional AC-DC electrical power converter electrically connected between the second electrical machine and the corresponding electrical distribution control unit. In this system configuration, the power may be taken only from the low pressure shaft. Such a take-off makes it possible to supply the AC aircraft network and to supply the HVDC aircraft network and / or the loads of the turbomachine via the unidirectional AC-DC electrical power converter and the PDMU.

[0025] In a fourth aspect of the system, each electrical distribution control unit may be intended to be further connected to a continuous electrical network main part of the aircraft.

[0026] Thus, the architecture makes it possible to supply local domestic loads of the turbomachine via the PDMU and to provide a possibility of taking or injecting power on the high pressure shaft via the local HVDC network.

[0027] In a fifth aspect of the system, each electrical distribution control unit may lack an electrical connection to a main DC electrical network of the aircraft.

[0028] This configuration thus makes it possible to supply the domestic loads of the turbomachine via the local HVDC network, and to take or inject power onto the high pressure shaft via the local HVDC network, and to start the turbomachine by injecting power onto the high pressure shaft via the AC electrical network coming from the aircraft.

[0029] In a sixth aspect of the system, the first electrical machine may be a wound rotor synchronous electrical machine, such as a variable frequency generator (VFG) for example composed of 2 independent stator windings directly distributed to the aircraft network.

[0030] This configuration makes it possible to dispense with all power converters and the HVDC network.

[0031] According to another aspect of the invention, there is provided a turbomachine of an aircraft comprising a high pressure shaft and a low pressure shaft and a system for supplying electrical equipment of an aircraft as defined previously.

[0032] According to yet another aspect of the invention, an aircraft is proposed comprising at least one turbomachine as defined previously. Brief description of the drawings

[0033] The invention will be better understood from the following reading, for informational but non-limiting purposes, with reference to the appended drawings in which:

[0034] [Fig-1] [Fig.l] shows a turbomachine of an aircraft equipped with a system power supply for electrical equipment of the aircraft according to a first embodiment of the invention.

[0035] [Fig.2] [Fig.2] illustrates a turbomachine of an aircraft provided with a system for supplying electrical equipment of the aircraft according to a second embodiment of the invention.

[0036] [Fig.3] [Fig.3] illustrates a turbomachine of an aircraft provided with a system for supplying electrical equipment of the aircraft according to a third embodiment of the invention.

[0037] [Fig.4] [Fig.4] illustrates a turbomachine of an aircraft equipped with a system for supplying electrical equipment of the aircraft according to a fourth embodiment lization of the invention. Description of the embodiments

[0038] In [Fig.l] is illustrated a turbomachine of an aircraft provided with a system for supplying electrical equipment of the aircraft according to a first embodiment of the invention.

[0039] 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 further comprises a high pressure compressor 5 and a high pressure turbine 6 coupled to a high pressure shaft 7.

[0040] The turbomachine 1 further comprises an electrical power supply system 8 which is a hybrid electric system for powering electrical equipment. The power supply system 8 comprises a first electrical machine 9 and a second electrical machine 10.

[0041] 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 permanent magnet synchronous electrical machine.

[0042] 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 wound rotor synchronous electrical machine comprising an electrical output directly connected to a main AC electrical network 11 of the aircraft.

[0043] For each first stator winding of the first electrical machine 9, the power supply system 8 comprises an electrical distribution control unit 12 (PDMU) and a bidirectional DC-AC electrical power converter 13 electrically connected between the first stator winding and the electrical distribution control unit 12, the electrical distribution unit 12 being connected at the output to a high-voltage DC electrical network 16 of the aircraft and to electrical loads 14 of the turbomachine 1.

[0044] 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 electrically connected between the second electrical machine 10 and the corresponding electrical distribution control unit 12.

[0045] Each electrical distribution control unit 12 is further connected to a main direct current electrical network 16 (HVDC) of the aircraft.

[0046] Thus, the architecture makes it possible to supply local domestic loads 14 from the tur- bomachine 1 via PDMU 12 and provide a possibility to take or inject power on high pressure shaft 7 by the local HVDC network.

[0047] In [Fig.2] is illustrated a turbomachine of an aircraft equipped with a system power supply of electrical equipment of the aircraft according to a second embodiment of the invention.

[0048] The second embodiment illustrated in [Fig.2] differs from the first embodiment illustrated in [Fig.l] in that each electrical distribution control unit 12 lacks an electrical connection to a main direct current (HVDC) electrical network of the aircraft.

[0049] This configuration thus makes it possible to supply the domestic loads of the turbomachine via the local HVDC network, and to take or inject power onto the high pressure shaft via the local HVDC network, and to start the turbomachine by injecting power onto the high pressure shaft via the AC network coming from the aircraft.

[0050] In [Fig.3] is illustrated a turbomachine of an aircraft equipped with a system power supply of electrical equipment of the aircraft according to a third embodiment of the invention.

[0051] The third embodiment differs from the first embodiment illustrated in [Fig.l] in that the second electrical machine 10 is devoid of electrical connection to the electrical distribution control unit 12.

[0052] In this third embodiment, the second electrical machine 10 coupled to the low pressure shaft 3 is thus solely dedicated to the main electrical generation of the aircraft in AC.

[0053] Alternatively, an HVDC network may be reconstituted on the aircraft side and interface with the local HVDC network of the turbomachine via the PDMU.

[0054] In [Fig.4] is illustrated a turbomachine of an aircraft equipped with a system power supply of electrical equipment of the aircraft according to a fourth embodiment of the invention.

[0055] The fourth embodiment differs from the first embodiment illustrated in [Fig.l] in that the first electrical machine 90 is a wound rotor synchronous electrical machine 90, such as a variable frequency generator (VFG) for example composed of 2 independent stator windings directly distributed to the aircraft network, and in that the electrical power system 8 is devoid of an electrical distribution control unit, a bidirectional AC-DC electrical power converter and a unidirectional AC-DC electrical power converter.

[0056] This configuration makes it possible to dispense with all power converters and the HVDC network.

[0057] The invention thus makes it possible to provide a hybrid electric system for supplying electrical equipment of an aircraft from a turbomachine making it possible to carry out the hybridization functions and the main generation of the aircraft with a known electrical network already in service with aircraft manufacturers in order to limit the risks and find a realistic interface for aircraft manufacturers between the hybrid electric network and the aircraft electrical network.

Claims

Claims

1. Hybrid electric system (8) for supplying 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 supply system (8) comprising: - a first electric 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 electric machine (10) intended to be mechanically connected to the low pressure shaft (3) of the turbomachine (1) and comprising two second independent stator windings, characterized in that the second electric machine (10) is a wound rotor synchronous electric machine comprising an electrical output intended to be directly connected to a main AC electrical network (11) of the aircraft.

2. System according to claim 1, wherein the first electrical machine (9) is a permanent magnet synchronous electrical machine, and the system (8) further comprises, for each first stator winding, an electrical distribution control unit (12) and a bidirectional DC-AC electrical power converter (13) electrically connected between the first stator winding and the electrical distribution control unit (12), said electrical distribution unit (12) being connected at output to a high voltage DC electrical network (HVDC) of the aircraft (16) and to electrical loads (14) of the turbomachine (1).

3. System (8) according to claim 2, wherein the second electrical machine (10) is devoid of electrical connection to an electrical distribution control unit (12).

4. System (8) according to claim 2, further comprising, for each second stator winding, a unidirectional AC-DC electrical power converter (15) electrically connected between the second electrical machine (10) and the corresponding electrical distribution control unit (12).

5. System (8) according to claim 4, in which each electrical distribution control unit (12) is intended to be further connected to a main continuous electrical network (16) of the aircraft.

6. System (8) according to claim 4, wherein each unit of electrical distribution control (12) is devoid of an electrical connection to a main DC electrical network of the aircraft.

7. The system (8) of claim 1, wherein the first electrical machine (90) is a wound rotor synchronous electrical machine (90).

8. Turbomachine (1) of an aircraft comprising a high pressure shaft (7) and a low pressure shaft (3) and an electrical equipment supply system (8) of an aircraft according to one of claims 1 to 7.

9. Aircraft comprising at least one turbomachine (1) according to claim 8.

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