Method for verifying the maximum available power of different components of an aircraft propulsion system.

Abstract

A method for verifying the maximum available power of components of an aircraft propulsion system comprising at least one first component sized to compensate for a failure of at least one second component of the propulsion system by providing maximum power to maintain the aircraft within a safe operating range. The method comprises the following steps for each of the first components: setting the first component to a speed substantially equal to a maximum power speed; adjusting the power supplied by the second component operating in synergy with the first component so that the first and second components contribute to providing the necessary power to the aircraft during the flight phase; determining the power supplied by the first component at its maximum power speed; and deducing from the determined power information relating to the maximum available power of the first component. (Figure from the abstract: Fig. 3)

Description

Title of the invention: Method for verifying the maximum available power of different components of an aircraft propulsion system.

[0001] The present invention relates to the propulsion system of aircraft and more particularly to a method of verifying the maximum available power of components of an aircraft propulsion chain.

[0002] BACKGROUND OF THE INVENTION

[0003] For example, a helicopter is an aircraft typically equipped with a main rotor driving a rotating wing to provide lift and propulsion. To drive the main rotor, it is known to equip the helicopter with a redundant propulsion system comprising two turbomachines operating in parallel. Each is designed to be oversized so that, in the event of a failure of the other turbomachine, it can provide sufficient power to allow the helicopter to continue its flight and land safely.

[0004] However, it is not possible to test the aircraft at such an operating regime called OEI (“One Engine Inoperative” or with a failed engine), since the valid turbomachine provides a power much greater than its rated power and is consequently damaged, so that it can no longer be used without heavy maintenance action.

[0005] It was therefore necessary to develop strategies to ensure the availability of maximum power at the OEI regime for each of the turbomachines.

[0006] It has been envisaged that, during a commercial flight, for example every 25 hours, a series of temperature and rotational speed measurements could be taken for each of the turbomachines in order to determine whether, from a thermodynamic point of view, each turbomachine would be capable of delivering maximum power at the OEI regime. This check, called EPC (Engine Power Check), is supplemented by specific maintenance operations aimed at detecting any latent failures of certain turbomachine components.

[0007] However, this strategy is not entirely satisfactory. Indeed, EPC control is performed at a power level significantly lower than the power required for takeoff, which increases uncertainty regarding the turbomachinery's ability to deliver maximum power at the OEI regime. Moreover, maintenance operations are particularly delicate and complicated to implement, requiring specific tools and qualified personnel, without which the risks of maintenance errors are high.

[0008] While document FR-A-1752849 describes a method for overcoming the problems described above, it only concerns the case where the aircraft propulsion system includes two thermal engines (such as turbomachines) operating in parallel.

[0009] However, among aircraft, there are, for example, so-called hybrid helicopters which are equipped with a propulsion system comprising a single internal combustion engine assisted during certain phases of flight by a brushless electric motor. The energy required to operate the electric motor is supplied by an auxiliary power unit (also called an "APU") or by an electric generator, which can also fail.

[0010] There are also aircraft equipped with an exclusively electric propulsion system which includes one or more electric motors powered by at least one battery.

[0011] It appears that the teaching of document FR-A-1752849 cannot be applied directly to aircraft equipped with a propulsion system having multiple redundancies of components (active or passive and standby), as may be the case with hybrid propulsion aircraft (thermal and electric engine) or exclusively electric propulsion aircraft.

[0012] SUBJECT OF THE INVENTION

[0013] The invention aims to provide a solution that at least partially remedies the aforementioned drawbacks. Summary of the invention

[0014] To this end, a method is proposed for verifying the maximum available power of components of an aircraft propulsion system comprising at least one first component sized to compensate for a failure of at least one second component of the propulsion system by providing maximum power to maintain the aircraft within a safe operating range, the method comprising the following steps for each of the first components:

[0015] - to set the first organ to a regime substantially equal to a power regime maximum;

[0016] - adjust a power supplied by the second organ operating in synergy with the first organ so that the first organ and the second organ contribute to providing the necessary power to the aircraft during the flight phase;

[0017] - determine a power supplied by the first component brought to power maximum;

[0018] - deduce from the determined power information relating to the power maximum available from the first organ.

[0019] Maximum power refers to the maximum power under specific operating conditions. There are several maximum power levels, such as maximum power at takeoff for different temperatures and pressures, maximum power in hover mode outside ground effect, etc. This method has the advantage of using the maximum power regime to ensure that each of the primary components can deliver maximum power at each regime or worst-case operating condition, particularly at regimes corresponding to particularly high power levels, such as during takeoff, in the OEI regime, and in hover mode outside ground effect. Indeed, at the maximum power regime, the power level is high enough to limit uncertainties regarding the component's ability to reach the expected power level.This ensures that aging, degradation, limitations, failures or latent errors in the primary components do not prevent full power from being achieved when needed.

[0020] According to a particular feature, the process further comprises the following steps: - determine a threshold power corresponding to a minimum power to be reached by the first component put into maximum power mode, and - compare the power supplied by the first component to the threshold power.

[0021] According to another particular feature, the first organs comprise an electrical power source, an organ for distributing the electrical power supplied by the electrical power source and an organ for transforming the power supplied by the distribution organ.

[0022] In particular, the process is automatically interrupted when at least one of the following conditions is met: - the power supplied by the first component brought to maximum power is less than or equal to a threshold power; - a parameter of the first component brought to maximum power is below or above a predetermined threshold; - a fault is detected on the first component brought to maximum takeoff power.

[0023] In particular, the method further includes the step of collecting data from the first organs during the other steps, and enriching a model allowing predictive analysis of maintenance action needs for each of said first organs.

[0024] The invention also relates to a computer program product comprising instructions for the execution of such a process when this program is executed by a processor.

[0025] The invention also relates to a control device comprising a computer configured to implement such a process.

[0026] The invention further relates to an assembly comprising at least two chains of components configured to operate in synergy and together provide the necessary power to an aircraft during a phase of flight, the assembly being characterized in that it includes such a control device.

[0027] The invention also relates to an aircraft comprising at least two chains of components configured to operate in synergy and together provide the power necessary to an aircraft during a phase of flight, the aircraft being characterized in that it includes a computer configured to implement such a method. Brief description of the drawings

[0028] The invention will be better understood in the light of the following description, which is purely illustrative and not limiting, and should be read in conjunction with the accompanying figures, among which:

[0029] [Fig-1] [Fig. 1] is a schematic view of a hybrid electric propulsion aircraft;

[0030] [Fig.2] [Fig.2] is a view partially and schematically representing the architecture of the hybrid electric propulsion of the aircraft illustrated in [Fig.1];

[0031] [Fig.3] [Fig.3] illustrates a method for verifying the maximum available power of different components of an aircraft according to an embodiment of the invention. DETAILED DESCRIPTION OF THE INVENTION

[0032] With reference to [Fig.1], a vertical take-off and landing aircraft 1 (also called "Vertical Take-Off and Landing aircraft" or "VTOL"), here of the drone type, includes a propulsion system comprising four propellers H1, H2, H3, H4 serving both for propulsion and for the lift of the aircraft 1. Each of the propellers H1, H2, H3, H4 is driven in rotation by a first electric motor ML1, M2.1, M3.1, M4.1 and a second electric motor M1.2, M2.2, M3.2, M4.2. The first engine M1.1, M2.1, M3.1, M4.1 is sized to provide the minimum power required for propulsion and lift of aircraft 1 in the event of a failure of the second engine M1.2, M2.2, M3.2, M4.2, and vice versa. Under normal operating conditions, half of the power required by the propeller H1, H2, H3, H4 is supplied by the first engine M1.1, M2.1, M3.1, M4.1, the other half being supplied by the second engine M1.2, M2.2, M3.2, M4.2 in order to minimize wear on the electrical and mechanical components of said first and second engines M1.1, M1.2, M2.1, M2.2, M3.1, M3.2, M4.1, M4.2. .

[0033] The propulsion system also includes a turbogenerator TG connected to a first electrical power distributor D10 supplying the first engines M1.1, M2.1, M3.1, M4.1, and a second electrical power distributor D20 supplying the second engines M1.2, M2.2, M3.2, and M4.2. A first battery, BATI, is also connected to the first distributor D10 and is sized to supplement the power supplied by the turbogenerator TG and compensate for its failure. Similarly, a second battery, BAT2, is connected to the second distributor D20 and is sized to supplement the power supplied by the turbogenerator TG and compensate for its failure.

[0034] Thus, the first motors M1.1, M2.1, M3.1, M4.1 and the second motors M1.2, M2.2, M3.2, M4.2 operate in pairs to rotate the propellers H1, H2, H3, H4. The first and second distributors D10, D20 operate in synergy to supply the first motors M1.1, M2.1, M3.1, M4.1 and the second motors M1.2, M2.2, M3.2, M4.2. The turbogenerator TG and the first battery BATI operate in synergy to supply the first distributor D10. The turbogenerator TG and the second battery BAT2 operate in synergy to supply the second distributor D20.

[0035] Four failure scenarios can be envisaged on the propulsion system of aircraft 1: - the first engine Ml.l, M2.1, M3.1, M4.1 or the second engine M1.2, M2.2, M3.2, M4.2 of a propeller Hl, H2, H3, H4 fails: the remaining active engine is then stressed at a power level much higher than its nominal power in order to compensate for the loss of the other engine (the stress on the first and second distributors D10, D20, the turbogenerator TG and the first and second batteries BATI, BAT2 is modified to a lesser extent); - the first battery BATI or the second battery BAT2 fails: one solution is for the turbogenerator TG to compensate for the failure of the battery BATI, BAT2 so as to provide all the electrical power necessary for the first and second engines Ml.l, M2.1, M3.1, M4.1, M1.2, M2.2, M3.2, M4.2 to satisfy the minimum safety requirements of aircraft 1 (the demand on the first and second distributors D10, D20 then remains balanced and the turbogenerator TG is demanded at a power level significantly higher than its normal operating power), another solution is to unbalance the demand on the first and second distributors D10, D20 in order to distribute the required power on the turbogenerator TG and on one of the first and second batteries BATI, BAT2; - The TG turbogenerator fails: the first and second batteries BATI, BAT2 must then supply all the electrical power generated necessary for the first and second engines Ml.1, M2.1, M3.1, M4.1, M1.2, M2.2, M3.2, M4.2 to meet the minimum safety requirements of aircraft 1, which puts them under a power level generally close to their maximum operating power; - the first distributor D10 or the second distributor D20 fails (for example due to a short circuit or the breakage of an electrical cable): all the motors Ml.l, M1.2, M2.1, M2.2, M3.1, M3.2, M4.1, M4.2 supplied by the remaining distribution chain D10, D20 will then operate at a power level much higher than their nominal power, as will the turbogenerator TG and the corresponding battery BATI, BAT2.

[0036] Fig. 2 schematically represents the architecture of the propulsion system of aircraft 1. For clarity, only the propellers H1, H2 are shown.

[0037] The TG turbogenerator delivers electrical power WelO to the first distributor D10 and electrical power We20 to the second distributor D20. Similarly, the first battery BATI delivers electrical power Well to the first distributor D10 and the second battery BAT2 delivers electrical power We21 to the second distributor D20.

[0038] The first distributor D10 in turn delivers an electrical power Wel.1 to the first motor Ml.1 of the propeller Hl and an electrical power We2.1 to the first motor M2.1 of the propeller H2. Similarly, the second distributor D20 delivers an electrical power Wel.2 to the second motor M1.2 of the propeller Hl and an electrical power We2.2 to the second motor M2.2 of the propeller H2.

[0039] The first motor M1.1 and the second motor M1.2 deliver mechanical power W1.1 and W1.2 respectively, both of which are used to rotate the propeller HL. Similarly, the first motor M2.1 and the second motor M2.2 deliver mechanical power W2.1 and W2.2 respectively, both of which are used to rotate the propeller H2. In normal operation, the mechanical power required to rotate the propellers H1 and H2 is distributed equally between the corresponding first motor M1.1 and M2.1 and the second motor M1.2 and M2.2.

[0040] The propeller H1, the first engine M1.1, the second engine M1.2, the first and second distributors D10, D20, the first and second batteries BATI, BAT2, and the turbogenerator TG each form a component of the first propulsion system of aircraft 1. It is understood that these "components" or "power components" play a direct role in the production and use of the power necessary for the flight of aircraft 1. Similarly, the propeller H2, the first engine M2.1, the second engine M2.2, the first and second distributors D10, D20, the first and second Batteries BATI, BAT2 and the turbogenerator TG each form an element of a second propulsion chain of aircraft 1. Thus, the first and second distributors D10, D20, the first and second batteries BATI, BAT2 and the turbogenerator TG are elements common to the first and second propulsion chains of aircraft 1.

[0041] Aircraft 1 is further equipped with a control device comprising a CAL computer configured to act: - on the first and second engines Ml.l, M1.2, M2.1, M2.2 via control instructions Cl.l, C1.2, C2.1, C2.2; - on the first and second distributors D10, D20 via CIO, C20 control instructions; and - on the TG turbogenerator via CL control instructions

[0042] It should be noted that the first and second distributors D10, D20 may include active electrical converters allowing the computer CAL to vary, via instruction commands CIO, C20, the power delivered by the first and second batteries BATI, BAT2 and / or the turbogenerator TG.

[0043] The CAL computer thus makes it possible to act on each of the components of the first and second propulsion chains of the aircraft 1 so as, in particular in the event of failure of one of the components, to put the component, which works in synergy with the failed component, at a maximum power regime and to adjust the power supplied by the other components so as to satisfy the minimum safety requirements of the aircraft 1.

[0044] The CAL computer is also configured to acquire measurements of various operating parameters for each of the components of the first and second propulsion chains. In particular, the CAL computer can acquire: - mechanical and thermal measurements MMLi, MMi 2, Mm2.i, Mm22 for each of the first and second motors Ml.l, M 1.2, M2.1, M2.2 including the effective rotational speed, the transmitted torque and the operating temperature (taken for example at the level of windings or power transistors of the first and second motors Ml.l, M1.2, M2.1, M2.2); - electrical measurements MEEi, MEL2, ME2 b ME22 for each of the first and second motors Ml.l, M1.2, M2.1, M2.2 including the electric current flowing in the winding of the first and second motors Ml.l, M 1.2, M2.1, M2.2, the voltage and / or supply current of said first and second motors Ml.l, M 1.2, M2.1, M2.2; - electrical measurements Mi0, M20 for each of the two distributors D10, D20 including the current IDi0, Id2o and the voltage VDio, VD2o passing through said distributors D10, D20; - mechanical and thermal measurements Mm[ and electrical MEi of the turbogen- nerator TG including the temperature and rotational speed of the turbine of the turbogenerator TG, the torque supplied by said turbine to drive the electricity generator of the turbogenerator TG, the current flowing in the winding of said generator...

[0045] According to the invention, the CAL computer is further configured to implement a method for verifying the maximum available power of the components of the aircraft's propulsion system 1, the verification method comprising the following steps ([Fig.3]): • for each of the first Ml.l, M2.1 engines: - set, via command C1.1, C2.1, the first motor Ml.l, M2.1 to a regime substantially equal to a maximum power regime PM allowing to satisfy the minimum safety requirements of aircraft 1 (step 10), preferably under a low electrical voltage so as to precipitate the latent failures which the first motor Ml.l, M2.1 could present by strongly stressing the electrical components constituting said first motor Ml.l, M2.1; - adjust, via the Cl.2, C2.2 control, the power Wml.2, Wm2.2 supplied by the second engine Ml.2, M2.2 operating in synergy with the first engine Ml.l, M2.1 so that the first engine Ml.l, M2.1 and the second engine Ml.2, M2.2 continue to supply the necessary power to aircraft 1 during the flight phase (step 20); - determine the power Wml.l, Wm2.1 supplied by the first engine Ml.l, M2.1 operating at maximum power PM (step 30); and - deduce from the power Wml.l, Wm2.1 information relating to the maximum available power of the first motor Ml.l, M2.1 (step 50); • for each of the second engines Ml.2, M2.2: - put, via command Cl.2, C2.2, the second motor Ml.2, M2.2 to a regime substantially equal to a maximum power regime PM allowing to satisfy the minimum safety requirements of aircraft 1 (step 10), preferably under a low electrical voltage so as to precipitate the latent failures which the second engine Ml.2, M2.2 could present by strongly stressing the electrical components constituting said second engine Ml.2, M2.2; - adjust, via the Cl.l, C2.1 control, the power Wml.l, Wm2.1 supplied by the first engine Ml.l, M2.1 operating in synergy with the second engine Ml.2, M2.2 so that the first engine Ml.l, M2.1 and the second engine Ml.2, M2.2 continue to supply the necessary power to aircraft 1 during the flight phase (step 20); - determine the power Wml.2, Wm2.2 supplied by the second motor Ml.2, M2.2 set to maximum power mode to be verified (step 30); and deduce from the power Wml.2, Wm2.2 information relating to the maximum available power of the second motor Ml.2, M2.2 (step 50); for the first BATI battery: set, via the CIO command, the first BATI battery to a speed substantially equal to a maximum power PM speed allowing to satisfy the safety minimums of aircraft 1 (step 10); adjust, via the Cl control, the WelO power supplied by the TG turbogenerator so that the first BATI battery and the TG turbogenerator continue to supply the necessary power to aircraft 1 during the flight phase (step 20); determine the Wel 1 power supplied by the first BATI battery set to maximum power PM (step 30); and deduce from the Wel 1 power information relating to the maximum available power of the first BATI battery (step 50); for the second battery BAT2: set, via command C20, the second battery BAT2 to a regime substantially equal to a maximum power regime PM allowing to satisfy the safety minimums of aircraft 1 (step 10); adjust, via the Cl control, the We20 power supplied by the TG turbogenerator so that the second battery BAT2 and the TG turbogenerator continue to supply the necessary power to aircraft 1 during the flight phase (stage 20); determine the We21 power supplied by the second battery BAT2 set to maximum power mode PM (step 30); deduce from the We21 power information relating to the maximum available power of the second battery BAT2 (step 50); for the TG turbogenerator: set, via the Cl control, the turbogenerator TG to a speed substantially equal to a maximum power speed PM allowing to satisfy the minimum safety requirements of aircraft 1 (step 10); adjust, via the CIO command, the Wel 1 power supplied by the first BATI battery (or, via the C20 command, the We21 power supplied by the second BAT2 battery) so that the TG turbogenerator and the first BATI battery (or the second BAT2 battery) continue to supply the necessary power to aircraft 1 during the flight phase (stage 20); determine the WelO power (or the We20 power) supplied by the TG turbogenerator set to maximum power PM (step 30); deduce information from the WelO power (or the We20 power) relating to the maximum available power of the TG turbogenerator (step 50); • for each of the two distributors D10, D20: - put, via the CIO, C20 commands, one of the distributors D10, D20 at a speed substantially equal to a maximum power speed PM allowing to satisfy the safety minimums of aircraft 1 (step 10); - adjust, via the CIO command, C20, the power Wel.2, We2.2, Wel.1, We2.1 supplied by the other of the distributors D10, D20 so that the distributors D10, D20 continue to supply the necessary power to aircraft 1 during the flight phase (step 20); - determine the power Wel.1, Wel2.1, Wel.2, Wel2.2, supplied by the distributor D10, D20 set to maximum power PM (step 30); and - deduce from the power Wel.l, We2.1, Wel.2, We2.2 information relating to the maximum available power of the distributor DI0, D20 set to maximum power PM (step 50).

[0046] It should be noted that the term "operating regime" here refers to the operating conditions of the component in question. This is the "rotational speed / torque" pair for the M1.1, M2.1, M3.1, M4.1, M1.2, M2.2, M3.2, and M4.2 engines, the "battery voltage / current" pair for the BATI and BAT2 batteries, and the "turbogenerator voltage / current" pair for the TG turbogenerator...

[0047] It will be understood that each of the components of the propulsion system is tested at a predetermined maximum power regime in order to ensure that the whole of the components does not have a latent failure, the said components being tested one after the other or in batches when the maximum power regime is identical.

[0048] The propulsive power Pprop required by aircraft 1 during the flight phase is, for example, entered into the control device, in particular the CAL computer, via a user interface connected to the control device. The user interface can also be configured to display status information about aircraft 1 to a pilot or operator.

[0049] The control device further includes a memory in which is recorded, for each of the components of the two propulsion chains, a predetermined threshold power Ps corresponding to a minimum power to be reached by the component put into maximum power mode PM in the event of failure of the one operating in synergy.

[0050] The CAL calculator then compares the power Wml.l, Wml.2, Wm2.1, Wm2.2, Wel.l, We2.1, Wel.2, We2.2, Well, WelO, We20, We21 supplied by the component set to maximum power PM to the threshold power Ps in order to ensure that the power supplied is greater than or equal to the threshold power Ps (step 40).

[0051] More specifically, if the power Wml.l, Wml.2, Wm2.1, Wm2.2, Wel.l, We2.1, If the power supplied by the component operating at maximum power (PM) is greater than or equal to the threshold power (Ps), the CAL is configured to instruct the user interface to inform the pilot or operator that the component operating at maximum power (PM) can supply the threshold power (Ps). The threshold power (Ps) then corresponds to the minimum guaranteed power in the event of a failure of the component whose power is being adjusted. Conversely, if the power supplied is less than the threshold power (Ps), the CAL is configured to instruct the user interface to inform the pilot or operator that the component operating at maximum power (PM) cannot supply the minimum guaranteed power and that maintenance must be scheduled for said component.

[0052] The CAL computer can determine the power Wml.l, Wml.2, Wm2.1, Wm2.2 supplied by each of the first and second motors Ml.l, M1.2, M2.1, M2.2 from the mechanical measurements MMEi, MME2, Mm2.i, Mm2 2 and / or from the electrical measurements MEEi, MEE2, Me2.i, ME2 2 acquired by the CAL computer.

[0053] The CAL computer can determine the power Wel.1, Wel.2, We2.1, We2.2 supplied by each of the distributors D10, D20 from the electrical measurements Mi0, M20 acquired by the CAL computer.

[0054] The CAL computer can determine the WelO, We20 power supplied by the TG turbo-generator from the MEi electrical measurements acquired by the CAL computer.

[0055] The CAL computer can also be configured to compare the measured temperature T of the Ml.l, M 1.2, M2.1, M2.2 engine set to maximum power PM with a predetermined threshold temperature Ts recorded in the memory of the control device, so as to ensure that the measured temperature is below the threshold temperature Ts (step 60).

[0056] More specifically, if the temperature T of the engine Ml.l, M1.2, M2.1, M2.2 at maximum power PM is less than the threshold temperature Ts, the CAL computer is configured to command the user interface to inform the pilot or operator that the temperature of the engine Ml.l, M1.2, M2.1, M2.2 at maximum power PM does not exceed the threshold temperature Ts, i.e. that there is no overheating of said engine Ml.l, M1.2, M2.1, M2.2 when the latter is at a maximum rotation speed corresponding to the maximum power PM.

[0057] Otherwise, that is, if the measured temperature T of the engine Ml.l, M1.2, M2.1, M2.2 operating at maximum power PM is greater than or equal to the threshold temperature Ts, the CAL computer is configured to instruct the user interface to inform the pilot or operator that the engine Ml.l, M1.2, M2.1, M2.2 operating at maximum power PM is overheating and that it is necessary to program a maintenance action on said engine Ml.l, M1.2, M2.1, M2.2.

[0058] The CAL computer can also be configured to compare the rotational speed V measured on the engine Ml.l, M 1.2, M2.1, M2.2 brought to maximum power setting PM to a threshold rotational speed Vs recorded in the memory of the control device, so as to ensure that the measured rotational speed V is greater than or equal to the threshold rotational speed Vs (step 60).

[0059] More specifically, if the rotational speed V of the engine Ml.l, M 1.2, M2.1, M2.2 set to maximum power PM is greater than or equal to the threshold rotational speed Vs, the CAL computer is configured to command the user interface to inform the pilot or operator that the engine Ml.l, M 1.2, M2.1, M2.2 set to maximum power PM can reach the threshold rotational speed Vs when the latter is at the maximum temperature corresponding to the maximum power PM.

[0060] Otherwise, i.e. if the rotation speed V of the engine Ml.l, M1.2, M2.1, M2.2 set to maximum power PM is less than the threshold rotation speed Vs, the CAL computer is configured to command the user interface to inform the pilot or operator that the rotation speed of the engine Ml.l, M1.2, M2.1, M2.2 set to maximum power PM is limited by the maximum temperature corresponding to the maximum power PM and that it is necessary to program a maintenance action on said engine Ml.l, M1.2, M2.1, M2.2.

[0061] The CAL computer can also be configured to compare the temperature Tg measured on the turbogenerator TG set to maximum power PM to a predetermined threshold temperature Tgs recorded in the memory of the control device, so as to ensure that the measured temperature Tg is lower than the threshold temperature Tgs (step 60).

[0062] More specifically, if the temperature Tg of the turbogenerator TG at maximum power PM is less than the threshold temperature Tgs, the CAL computer is configured to command the user interface to inform the pilot or operator that the temperature of the turbogenerator TG at maximum power PM does not exceed the threshold temperature Tgs, i.e. that there is no overheating of said turbogenerator TG when the turbine of the latter is at the maximum rotation speed corresponding to the maximum power PM.

[0063] Otherwise, that is, if the measured temperature Tg of the turbogenerator TG operating at maximum power PM is greater than or equal to the threshold temperature Tgs, the CAL computer is configured to instruct the user interface to inform the pilot or operator that the temperature of the turbogenerator TG operating at maximum power PM is overheating and that it is necessary to program a maintenance action on said TG turbogenerator.

[0064] To minimize the risk that the behavior of the propulsion system is impacted by the potential limitation or loss of one of the tested components, and lead to a significant impact on the workload of the crew, the verification process is automatically interrupted when at least one of the following conditions is met: - the power supplied by the component set to maximum power PM is less than or equal to the threshold power Ps; - a parameter of the component set to maximum takeoff power (PMD) is below or above a predetermined threshold (for example, the rotation speed of one of the engines Ml.l, M1.2, M2.1, M2.2 is below the threshold rotation speed Vs, the temperature of one of the engines Ml.l, Ml.2, M2.1, M2.2 is above the temperature Ts, the temperature of the turbogenerator is above the threshold temperature Tgs...); - A fault is detected in the component operating at maximum power PM.

[0065] Preferably, the verification procedure is performed during each flight for each component of the propulsion system. In other words, the maximum power that each component of the propulsion system can deliver is preferably verified during each flight.

[0066] Moreover, the verification procedure is preferably carried out during a flight phase in which the effect of a failure of one of the components would be minimal, for example during a cruise phase, near a diversion surface.

[0067] The verification process makes it possible to ensure that all the components of the propulsion system can deliver maximum power at each regime, in particular at regimes corresponding to particularly high powers such as takeoff (PMD regime) or OEI regime.

[0068] In particular, using the maximum power regime PM to check the maximum available power of each of the components is particularly advantageous insofar as at this regime, the level of power supplied by the component is not likely to damage it significantly if it does not have a latent failure.

[0069] Using the maximum power regime PM to check the maximum available power of each of the components also has the advantages of: - precipitating dormant or latent failures of the components; - reduce the duration of exposure to dormant faults, particularly when the verification process is carried out on each flight for each component; - to cause a possible failure of the component operating at maximum takeoff power (PMD) under flight conditions, particularly during cruise, under conditions where the consequences of such a failure are minimized (in the event of failure of one component, the other component operating in synergy will be less stressed during the cruise phase than during other flight phases, which limits the risks of cascading effects, i.e. the loss of one organ, then the other organ); - limit or even avoid maintenance operations and therefore human interventions on components which can themselves generate new risks and immobilize the aircraft; - can be supplemented by an EPC control; and - to allow the application of the verification process to all types of flights, and in particular commercial flights.

[0070] The operating parameters of the various components collected during the verification process can also be stored in the memory of the control device for analysis on the ground to determine whether the components can continue to be used. The results of these analyses make it possible, for example, to better guarantee the availability of maximum power from each component at each operating regime for subsequent flights.

[0071] Furthermore, the verification procedure also has the advantage of being able to be carried out on any type of flight (commercial or technical) and of not disrupting the latter in terms of speed, altitude...

[0072] It can be verified beforehand that the electrical power consumed by aircraft 1 is sufficient to bring the propulsion system components to or near their maximum power output. To do this, the following steps are carried out: - measure operating parameters of the propulsion system allowing to evaluate the power delivered by each of the components whose maximum available power is to be determined (the parameters measured are for example the rotation speed of the engines Ml.l, M1.2, M2.1, M2.2, the electric current flowing in the windings of said engines Ml.l, M 1.2, M2.1, M2.2, the electric current delivered by the batteries BATI, BAT2 or by the turbogenerator TG), the measurement period having to be long enough (several seconds, or even tens of seconds) to allow to have an estimate of the average power consumed and reject short transient variations; - determine the power consumed by each group of components from the measured parameters (the groups of components can for example be the engines Ml.l, M 1.2, M2.1, M2.2 of the same propeller Hl, H2, the battery and the turbogenerator supplying the same propulsion chain...); - compare the power consumed by each group of organs with a predetermined threshold power guaranteeing the group's capacity to perform the verification of maximum power availability under satisfactory operating and safety conditions; - allow verification of maximum power availability when the power consumed by each group of components is greater than the threshold power defined for each of them.

[0073] The verification process may also include an additional step of collecting data from the propulsion system components and enriching a model that allows for predictive analysis of maintenance needs for each of these propulsion system components. This analysis can, for example, be carried out by monitoring the level and trend of the measured parameters over time. Maintenance personnel can thus be informed of upcoming maintenance on one or more components, particularly when the level of one parameter degrades more rapidly than the others, while remaining within an acceptable range.

[0074] Of course, the invention is not limited to the embodiment described but encompasses any variant falling within the scope of the invention as defined by the claims.

[0075] The aircraft propulsion system architecture may include both electric motors and internal combustion engines that can run on various fuels (kerosene, hydrogen, etc.). In other words, one or more of the engines M1.1, M1.2, M2.1, M2.2 of aircraft 1 may, for example, be replaced by internal combustion engines.

[0076] The aircraft propulsion system architecture may include batteries, fuel cells, or other sources of electrical power. In other words, the turbogenerator and / or one of the batteries may, for example, be replaced by a fuel cell.

[0077] The method applies to any architecture combining different sources of energy, power distribution and transformation.

[0078] The method applies to all types of aircraft: VTOL (Vertical Take-Off and Landing aircraft), STOL (Short Take-Off and Landing aircraft), or conventional aircraft...

[0079] The components of the propulsion system can be tested individually or in groups.

[0080] The maximum power regime PM can, for example, correspond to the maximum power regime at takeoff.

[0081] The aircraft may be fixed-wing or rotary-wing, manned or unmanned, with vertical, short, or long takeoff...

Claims

Demands

1. Method for verifying the maximum available power of components of an aircraft propulsion system (1) comprising first components (Ml.1, M2.1, D10, TG, BATI) sized to compensate for a failure of second components of the propulsion system (Ml.2, M2.2, D20, TG, BAT2) by providing maximum power to maintain the aircraft in a safe operating range, the first components comprising an electrical power source (TG, BATI, BAT2), a distribution component (D10, D20) of the electrical power supplied by the electrical power source and a transformation component (Ml.1, Ml.2, M2.1, M2.2) of the power supplied by the distribution element, the process comprising the following steps for each of the first elements: - set the first element to a regime substantially equal to a maximum power (PM) regime; - adjust a power supplied by the second element operating in synergy with the first element so that the first element and the second element contribute to supplying the power required by the aircraft during the flight phase; - determine a power supplied by the first element set to the maximum power regime; - deduce from the determined power an information relating to the maximum available power of the first element.

2. Method according to claim 1, comprising the following additional steps: - determine a threshold power (Ps) corresponding to a minimum power to be reached by the first component brought to maximum power (PM), and - compare the power supplied by the first component to the threshold power (Ps).

3. A method according to any one of the preceding claims, automatically interrupted when at least one of the following conditions is met: - the power supplied by the first component (Ml.1, Ml.2, M2.1, M2.2, D10, D20, TG, BATI, BAT2) set to maximum power (PM) is less than or equal to a threshold power (Ps), - a parameter of the first component set to maximum power (PM) is less than or greater than a predetermined threshold (Vs, Ts, Tgs), - a fault is detected on the first component set to maximum takeoff power (PMD).

4. A method according to any one of the preceding claims, further comprising the step of collecting data from the first components (Ml.1, M2.1, D10, TG, BATI) during the other steps, and enriching a model enabling predictive analysis of maintenance action needs for each of said first components.

5. Product computer program comprising instructions for carrying out the process according to any one of claims 1 to 4 when this program is executed by a processor.

6. Control device comprising a computer (CAL) configured to implement the method according to any one of claims 1 QA

7. d H-. Assembly comprising at least two chains of components configured to operate in synergy and together provide power necessary to an aircraft (1) during a phase of flight, the aircraft being characterized in that it comprises a control device according to claim 6.

8. Aircraft (1) comprising at least two chains of components configured to operate in synergy and together provide power necessary to an aircraft during a phase of flight, the aircraft being characterized in that it comprises a computer (CAL) configured to implement the method according to claim 1.

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