Method for monitoring the operation of a rotorcraft power plant and associated rotorcraft

Abstract

The present invention relates to a method for monitoring an operation of a power plant (2) of a rotorcraft (1) and a rotorcraft (1) comprising at least one lift rotor (5) driven in rotation by the power plant (2) comprising at least one engine (3, 4). Such a rotorcraft (1) comprises a monitoring system (6) comprising at least one failure sensor (23, 24) detecting a regulation failure (PAN) affecting a regulation system (13, 14) of said at least one engine (3, 4), a speed sensor (7) measuring a current rotation speed (NR) of said at least one lift rotor (5) and a controller (8) determining a compatibility condition (COMP) or an incompatibility condition (INCOMP) between a current flight phase and the rotation speed (NR). Abstract figure: figure 1

Description

Title of the invention: Method for monitoring the operation of a rotorcraft power plant and associated rotorcraft

[0001] The present invention relates to a method for monitoring the operation of a rotorcraft power plant.

[0002] A rotorcraft has at least one rotor participating in its lift, or even its propulsion.

[0003] For example, a rotorcraft may comprise a main rotor participating in its lift and propulsion, this main rotor having blades with collectively and cyclically variable pitch. In addition, the rotorcraft may comprise a system participating in the control of the yaw movement, such as another main rotor or a tail rotor for example. The pitch control system can be activated for example by the pilot or the crew. Alternatively, the system can be activated by an automatic control system, called an autopilot.

[0004] To rotate the rotor(s), the rotorcraft comprises a power plant, possibly multi-engine, and a power transmission chain going from the engines to one or more rotors.

[0005] For example, two motors are connected to a power transmission box, this power transmission box rotating the rotor(s). The power transmission box then comprises one mechanical input per motor, the mechanical inputs being engaged with a combiner, the combiner driving one mechanical output of the power transmission box per rotor. The motor inputs, the combiner, and the mechanical outputs may comprise at least one pinion or toothed wheel, at least one shaft, at least one speed reduction stage, etc.

[0006] The engines may be heat engines having an output shaft driven by the combustion of fuel. For example, at least one engine may be a turboshaft engine having a gas generator and a free working turbine connected to the output shaft.

[0007] Each engine can be controlled by a regulation system known by the acronym FADEC and the English expression "Full Authority Digital Engine Control". Such a regulation system comprises an engine computer in communication with sensors measuring values of operating parameters of the controlled engine, or even of the rotorcraft, such as a rotation speed of a gas generator or a free turbine, an internal temperature, a rotation speed of a rotor of the rotorcraft.

[0008] This engine computer is then configured to control a fuel metering device supplying the associated engine in order, for example, to stabilize a rotation speed of the output shaft at a set speed, regardless of the power consumed by that output shaft. This power varies depending on a pilot's actions on flight controls, particularly on controls that collectively and / or cyclically vary the pitch of the main rotor blades on a rotorcraft. Alternatively, these flight controls may be actuated by an autopilot.

[0009] Regulation as a function of the rotation speed of the output shafts of the engines then ensures a permanent match between the power produced jointly by the output shafts of these engines and the power consumed by the rotor(s), so that this or these rotor(s) rotate at a nominal rotation speed compatible with its or their function of generating lift for the rotorcraft.

[0010] Nevertheless, the engine computer is configured to prevent the power developed by an engine from causing the engine or the power transmission chain to exceed an acceptable mechanical limit. Usually, various operating regimes associated with specific limits and usage durations are then defined. For each operating regime, these limits may include, for example, a torque limit of an engine torque delivered by an engine, a limit of an internal temperature of an engine, a limit of rotational speed of a shaft of an engine, a torque limit of at least one member of the power transmission chain, etc.

[0011] On a multi-engine aircraft, there are known operating regimes called "AEO", for the English expression "All Engines Operative" usable for certain durations when all the engines are operating normally, and operating regimes called "OEI", for the English expression "One Engine Inoperative" usable for certain durations when one of the engines is out of order. The use of certain operating regimes can induce a maintenance action on the engines, or even on the power transmission chain.

[0012] An engine computer then controls a fuel metering control system so as to modulate the power of the associated engine. The fuel metering control system is thus configured to cause the rotational speed of the output shaft to tend towards a set value, while avoiding exceeding a limit.

[0013] Furthermore, in the event of a failure or breakdown of the engine control system, the fuel flow supplying the engine concerned is frozen at the last flow value used. A pilot is informed of this and can then either shut down the engine concerned for safety reasons or continue the flight as is, i.e. with an unregulated engine.

[0014] More generally, managing a lack of regulation of one of the engines of a rotorcraft can be complex to implement in order to pilot such a rotorcraft safely.

[0015] The present invention therefore aims to propose a method for limiting the workload of a crew in the presence of a failure of an engine regulation system on a rotorcraft or in the absence of regulation of one of the engines. Furthermore, this method and the associated rotorcraft can allow a pilot to carry out piloting maneuvers in complete safety, without risking degrading the mechanical power transmission chain or a rotor of the rotorcraft.

[0016] The invention thus relates to a method for monitoring the operation of a rotorcraft power plant, the power plant comprising at least one engine, the rotorcraft comprising at least one lift rotor driven in rotation by the power plant.

[0017] According to the invention, such a monitoring method is remarkable in that it comprises the following steps: - detection of a regulation failure affecting a regulation system of said at least one engine, - detection of a current rotation speed NR of said at least one lift rotor, - determination of a compatibility condition or an incompatibility condition between a current flight phase and the NR rotation speed, - in a first operating mode, in the presence of the compatibility condition and the regulation failure, generation of a first alert simultaneously representing the regulation failure and the compatibility condition, and - in a second operating mode, in the presence of the incompatibility condition and the regulation failure, generation of a second alert simultaneously representing the regulation failure and the incompatibility condition.

[0018] Furthermore, during a control failure of an engine, the fuel metering device supplying fuel to a combustion chamber of the associated engine then remains blocked in a current position and continues to deliver a constant fuel flow. Depending on the current flight phase, and in particular a change in pitch control of the rotor blades, the rotation speed NR of the rotor can then change without being controlled and go outside a nominal operating range.

[0019] Furthermore, the current flight phase may be identified by the monitoring method or identified by another method and transmitted to enable the determination of the compatibility condition or alternatively the incompatibility condition to be carried out.

[0020] In the first operating mode, such a monitoring method allows a pilot to be able to continue a flight phase despite the regulation failure of at least one engine, as long as the rotation speed NR is maintained within a nominal operating range for the current flight phase.

[0021] The first alert is thus generated solely for information purposes and the pilot can continue his mission without any particular action to be taken.

[0022] In the second operating mode, if the regulation failure of at least one engine is still present and the rotation speed NR goes outside the nominal operating range for the current flight phase, then the second alert can allow the pilot to carry out a predetermined emergency procedure. For example, the pilot can carry out an action allowing the rotation speed NR to return to its nominal range or to shut down the engine affected by the regulation failure.

[0023] Such a second alert is distinct from the first alert to allow a pilot to differentiate them.

[0024] Furthermore, each alert may take the form of a visual alarm, for example by emitting light with a light-emitting diode or equivalent or by displaying one or more characters on a screen, an audible alarm, by means of a loudspeaker, and / or a haptic alarm, for example by means of a vibrating unit vibrating an organ held or worn by an individual.

[0025] According to a first exemplary embodiment of the invention, the compatibility condition can be determined when the current rotation speed NR belongs to an interval of values defined by a rotation speed setpoint less a first margin (n) and the rotation speed setpoint plus a second margin (m) and alternatively the incompatibility condition can be determined when the current rotation speed NR is excluded from this interval of values.

[0026] In other words, in this case, the compatibility condition and the incompatibility condition are functions of a current rotation speed setpoint, a first margin (n) and a second margin (m). The first margin (n) and the second margin (m) may for example consist of a percentage of the rotation speed setpoint.

[0027] In practice, the rotation speed instruction can be variable depending on the current flight phase.

[0028] Such a rotation speed instruction can be generated by a control computer of an avionics system of the rotorcraft as a function in particular of atmospheric conditions of the flight phase and of the control instructions allowing the pilot to control the rotorcraft.

[0029] Similarly, the first margin (n) and the second margin (m) can be variable depending on the current flight phase.

[0030] Indeed, depending on the current flight phase, the first and second margins (n) and (m) can be modified and thus make it possible to adapt the limits of the nominal operating range corresponding to the acceptable values for the rotation speed NR of the rotor.

[0031] According to a second exemplary embodiment of the invention, the compatibility condition can be determined when the current rotation speed NR belongs to a predetermined range of acceptable values allowing the current flight phase to continue and alternatively the incompatibility condition can be determined when the current rotation speed NR is excluded from the predetermined range of acceptable values allowing the current flight phase to continue.

[0032] In this case, the limits of the nominal operating range corresponding to the acceptable values for the rotation speed NR of the rotor are fixed and predefined for example by ground tests, flight tests or simulations.

[0033] In this case, the predetermined range may include an upper limit defined so that a tangential speed at the blade tip of a blade of said at least one lift rotor is kept lower than the speed of sound.

[0034] Such an upper limit can thus make it possible to guarantee the integrity of the lift rotor. Alternatively or additionally, this upper limit can be defined in order to provide the rotorcraft with maximum performance during a so-called stationary or low-speed flight phase and / or in order to limit the acoustic footprint generated by the rotation of the rotor in a cabin of the rotorcraft.

[0035] Similarly, the predetermined range may include a lower bound defined to provide a minimum thrust allowing the rotorcraft to fly at a constant altitude at a cruising forward speed.

[0036] The lower limit can be chosen to reduce the acoustic footprint of a rotorcraft on the external environment. Indeed, such an acoustic footprint is directly linked to the rotation speed NR of the rotor and increases when the rotation speed NR increases.

[0037] Alternatively or additionally, this lower limit can be defined in order to avoid an excessively significant drop in altitude of the rotorcraft in the event of an engine failure and before being able to reach a rotor rotation speed allowing an autorotation flight phase to be carried out.

[0038] In practice, a transition from the first operating mode to the second operating mode may be irreversible.

[0039] Such irreversibility is in fact an additional protection. It helps to contribute to flight safety and to minimize the risks of confusion for the pilot.

[0040] Advantageously, the generation of the first alert may include a first display on a display of at least one item of information represented with a first predetermined color.

[0041] This or these pieces of information may, for example, include an alphanumeric message or a geometric shape such as a banner, a rectangle, a diamond, a circle, a line, etc.

[0042] The display may for example comprise a screen on a dashboard of the rotorcraft or a so-called "head-up" display device which may in particular be worn by the pilot and be integrated into glasses or a helmet screen.

[0043] The first predetermined color may be, for example, amber or orange.

[0044] Similarly, the generation of the second alert may include a second display on the display of said at least one item of information represented with a second predetermined color different from the first predetermined color.

[0045] This or these pieces of information may, for example, include an alphanumeric message or a geometric shape such as a banner, a rectangle, a diamond, a circle, a line, etc.

[0046] The message and / or the geometric shape of this second alert may be the same as those of the first alert, but the second predetermined color may be, for example, red and therefore distinct from an amber or orange color.

[0047] Furthermore, said at least one motor may for example comprise a first motor and a second motor. In this case, when switching to the second operating mode, the method may implement other additional steps aimed at bringing the rotation speed NR of the rotor back into its nominal operating range.

[0048] Thus, according to a first variant, the regulation failure affecting a first regulation system of the first engine, when the second alert is generated, the method can comprise a command to stop the first engine.

[0049] Such a command to stop the first engine may for example consist of closing a fuel supply solenoid valve of the first engine. Once the solenoid valve is closed, a first fuel metering device connected to the combustion chamber of the first engine can no longer supply fuel and the first engine stops. But the rotor remains rotated by the second engine, the regulation system of which is not affected by a failure.

[0050] According to a second variant, the regulation failure affecting a first regulation system of the first engine, when the second alert is generated, the method may comprise a control of a reversible transmission device to prevent the transmission of engine torque from the first engine to a power transmission chain.

[0051] Such control of a reversible transmission device of the first motor may for example consist of disengaging a clutch system or a controlled freewheel arranged between an output shaft of the first motor and an input shaft of a power transmission box. In this case, the first motor does not stop, but it no longer supplies engine torque to the rotor via the power transmission chain.

[0052] The present invention also relates to a rotorcraft comprising at least one rotor of lift driven in rotation by a power plant, the power plant comprising at least one motor.

[0053] According to the invention, such a rotorcraft is remarkable in that it comprises a surveillance system comprising: - at least one failure sensor detecting a regulation failure affecting a regulation system of said at least one engine, - a speed sensor measuring a current rotation speed NR of said at least one lift rotor, and - a controller determining a compatibility condition or an incompatibility condition between a current flight phase and the NR rotation speed, the controller generating, in a first operating mode, in the presence of the compatibility condition and the regulation failure, a first alert simultaneously representing the regulation failure and the compatibility condition, and the controller generating, in a second operating mode, in the presence of the incompatibility condition and the regulation failure, a second alert simultaneously representing the regulation failure and the incompatibility condition.

[0054] Furthermore, such a failure sensor can be integrated into the regulation system of said at least one engine, which is connected by wire or wireless means to the controller. The failure sensor thus directly transmits to the controller the information of a failure of the regulation system concerned.

[0055] Similarly, the speed sensor can be integrated into an avionics system which is connected by wire or wireless means to the controller. Such an avionics system also transmits information representative of the flight conditions to deduce a current flight phase.

[0056] From the information received, the controller deduces the compatibility condition or the incompatibility condition between the current flight phase and the rotation speed NR.

[0057] When receiving the information about the regulation failure, the controller can then determine whether it must implement the first operating mode or the second operating mode and thus generate the corresponding alert.

[0058] The invention and its advantages will appear in more detail in the context of the description which follows with examples given for illustrative purposes with reference to the appended figures which represent:

[0059] [Fig.l], a diagram of a rotorcraft in side view equipped with a surveillance system according to the invention,

[0060] [Fig.2], a flowchart illustrating a monitoring method according to the invention,

[0061] [Fig.3], a flowchart illustrating a first alternative of the monitoring method according to the invention, and

[0062] [Fig.4], a flowchart illustrating a second alternative of the monitoring method according to the invention.

[0063] Elements present in several distinct figures are assigned a single reference.

[0064] As already mentioned, the invention relates to a rotorcraft equipped with a system for monitoring an engine installation and the associated monitoring method.

[0065] As shown in [Fig.l], such a rotorcraft 1 comprises a power plant 2 provided with at least one engine 3, 4 allowing the rotational drive of at least one lift rotor 5 via a power transmission chain 11. Such a power transmission chain 11 may in particular comprise a power transmission box 9 having at least one input shaft connected respectively to an output shaft of each engine 3, 4.

[0066] Such a power transmission box 9 then comprises a mechanical output intended to drive in rotation a rotor mast integral in rotation with said at least one lift rotor 5.

[0067] Furthermore, the flow rate of fuel supplied to said at least one engine 3, 4 is initially regulated by a dedicated regulation system 13, 14 connected to a fuel metering device not shown.

[0068] The monitoring system 6 then comprises at least one failure sensor 23, 24 capable of detecting a PAN regulation failure affecting the regulation system 13, 14 of said at least one engine 3, 4.

[0069] The monitoring system 6 also comprises a speed sensor 7, for example arranged at the level of the rotor mast, and capable of measuring a current rotation speed NR of said at least one lift rotor 5.

[0070] By fault sensor 23, 24 and speed sensor 7, we mean here physical sensors capable of directly measuring the parameter in question, but also a system which may comprise one or more physical sensor(s) as well as signal processing means making it possible to provide an estimate of the parameter from the measurements provided by these physical sensors. Similarly, by measurement of this parameter we will designate both a raw measurement from a physical sensor and a measurement obtained by more or less complex signal processing from one or more raw measurements.

[0071] The monitoring system 6 then comprises a controller 8 connected by wire or wireless means to the speed sensor 7 and a set of other sensors not shown, for example from an avionics system of the rotorcraft 1, configured to enable a current flight phase of the rotorcraft 1 to be identified.

[0072] The set of sensors configured to identify a current flight phase may notably include, but is not limited to, an air data computer ADC, a heading and attitude reference system AHRS, a radar altimeter Rad Alt and a global positioning system by GPS satellites.

[0073] The air data computer ADC provides the monitoring system 6 in the non-limiting embodiment described, but not limited to, the barometric altitude, the speed of movement of the rotorcraft 1 relative to the air and the gross vertical speed of the rotorcraft 1.

[0074] The AHRS attitude and heading reference system provides the monitoring system 6 in the described non-limiting embodiment, but not limited to, attitude and heading, acceleration and descent rate information.

[0075] With the input from the ADC air data computer, the AHRS heading and attitude reference system provides an instantaneous vertical speed.

[0076] With input from a GPS global positioning system and / or an FMS flight avionics system, the AHRS attitude and heading reference system provides a combined navigation position and ground speed in each direction.

[0077] The Rad Alt radar altimeter provides the height of rotorcraft 1 above the ground and water.

[0078] The controller 8 thus makes it possible to determine a compatibility condition COMP or an INCOMP incompatibility condition between a current flight phase and the NR rotation speed.

[0079] Furthermore, the controller 8 may comprise, for example, at least one processor and at least one memory, at least one integrated circuit, at least one programmable system, at least one logic circuit, these examples not limiting the scope given to the expression “controller”. The term processor may also designate a central processing unit known by the acronym CPU, a graphics processing unit GPU, a digital unit known by the acronym DSP, a microcontroller, etc.

[0080] Furthermore, the monitoring system 6 may also comprise a display 10 making it possible to display to a pilot of the rotorcraft 1 two separate alert items of information when a PAN regulation failure is detected.

[0081] Furthermore, according to figures 2 to 4, the invention also relates to a method 30, 40 for monitoring the operation of the power plant 2 driving the lift rotor(s) 5 in rotation.

[0082] Furthermore, instructions or a computer program may be stored in a memory of the monitoring system 6. The monitoring system 6 may then execute these instructions or this program to implement the monitoring control method 30, 40.

[0083] Such a monitoring method 30, 40 thus comprises a detection 31, 41 of a PAN regulation failure affecting a regulation system 13, 14 of said at least one engine 3, 4.

[0084] The monitoring method 30, 40 also comprises a detection 32, 42 of the current rotation speed NR of said at least one lift rotor 5 and a determination 33, 43 of a compatibility condition COMP or an incompatibility condition INCOMP between a current flight phase and the rotation speed NR. Consequently, the compatibility condition COMP and the incompatibility condition INCOMP are two opposing conditions which cannot be verified simultaneously.

[0085] Therefore, in a first operating mode M0D1, in the presence of the compatibility condition COMP and the regulation failure PAN, the monitoring method 30, 40 comprises a generation 34, 44 of a first alert simultaneously representing the regulation failure PAN and the compatibility condition COMP.

[0086] Alternatively, and in a second operating mode M0D2, in the presence of the incompatibility condition INCOMP and the PAN regulation failure, the monitoring method 30, 40 comprises a generation 35, 45 of a second alert simultaneously representing the PAN regulation failure and the incompatibility condition INCOMP.

[0087] Furthermore, the transition from the first operating mode M0D1 to the second operating mode M0D2 may be irreversible for flight safety reasons.

[0088] The steps 34, 44 of generating a first alert and 35, 45 of generating a second alert are thus implemented by the controller 8 which is capable of generating at least two alerts distinct from one another and thus informing the pilot of the rotorcraft 1 of a current safety level allowing or prohibiting the continuation of a mission during a PAN regulation failure.

[0089] Furthermore, different alternatives make it possible to determine the conditions of COMP compatibility and INCOMP incompatibility.

[0090] According to a first alternative of the monitoring method 30 represented in [Fig.3], the compatibility condition COMP is for example determined when the current rotation speed NR belongs to an interval of values defined by a rotation speed setpoint NRcons minus a first margin n and the rotation speed setpoint NRcons plus a second margin m and alternatively the incompatibility condition INCOMP is determined when the current rotation speed NR is excluded from this interval of values.

[0091] In practice, such a rotation speed instruction NRcons can be variable depending on the current flight phase identified by the controller 8.

[0092] Similarly, the first margin n and the second margin m can also be variable depending on the current flight phase.

[0093] Furthermore, the generation 34 of the first alert may include a first display 341 on the display 10 of at least one item of information represented with a first predetermined color.

[0094] The generation 35 of the second alert may include a second display 351 on the display 10 of said at least one item of information represented with a second predetermined color different from the first predetermined color.

[0095] For example, the first predetermined color may be amber and the second predetermined color may be red.

[0096] Said at least one item of information may for example include an alphanumeric message “FAIL” indicating a PAN regulation failure and being supplemented with the regulation system concerned 13 or 14 “FADEC1” or “FADEC2”.

[0097] Furthermore, when the rotorcraft 1 comprises a first engine 3 and a second engine 4, and the PAN regulation failure affects the first regulation system 13 of the first engine 3, the method 30 may comprise a command 36 for a shutdown of the first engine 3.

[0098] Such a command 36 can then be implemented after the generation 35 of the second alert, for example automatically after a predetermined time interval or manually by a pilot of the rotorcraft 1.

[0099] According to a second alternative of the monitoring method 40 represented in [Fig.4], the compatibility condition COMP can be determined when the current rotation speed NR belongs to a predetermined range of acceptable values allowing the current flight phase to continue and alternatively the incompatibility condition INCOMP can be determined when the current rotation speed NR is excluded from this predetermined range of acceptable values.

[0100] For example, the predetermined range may include an upper limit defined so that a tangential speed at the blade tip of a blade 12 of said at least one lift rotor 5 is kept lower than the speed of sound.

[0101] Similarly, the predetermined range may include a lower bound defined to provide a minimum thrust allowing the rotorcraft 1 to perform flight at a constant altitude at a cruising forward speed.

[0102] Furthermore, the generation 44 of the first alert may include a first display 441 on the display 10 of at least one item of information represented with a first predetermined color.

[0103] The generation 45 of the second alert may include a second display 451 on the display 10 of said at least one item of information represented with a second predetermined color different from the first predetermined color.

[0104] Said at least one piece of information may for example include a luminous geometric shape such as a rectangle or an amber-colored band to represent the first alert, and alternatively, the same shape of rectangle or red banner to represent the second alert.

[0105] Furthermore, such a luminous geometric shape can be displayed in the background of a digital indicator displaying a current value of the rotation speed NR.

[0106] Furthermore, when the rotorcraft 1 comprises a first engine 3 and a second engine 4, and the PAN regulation failure affects the first regulation system 13 of the first engine 3, the method 40 may comprise a control 46 of a reversible transmission device to prevent the transmission of engine torque from the first engine 3 to the power transmission chain 11.

[0107] Such a command 46 can then be implemented after the generation 45 of the second alert, for example automatically after a predetermined time interval or manually by a pilot of the rotorcraft 1.

[0108] Naturally, the present invention is subject to numerous variations as to its implementation. Although several embodiments have been described, it is understood that it is not conceivable to exhaustively identify all possible modes. It is of course conceivable to replace a means described by an equivalent means without departing from the scope of the present invention and the claims.

Claims

Claims

1. Method for monitoring (30, 40) an operation of a power plant (2) of a rotorcraft (1), said power plant (2) comprising at least one engine (3, 4), said rotorcraft (1) comprising at least one lift rotor (5) driven in rotation by said power plant (2), characterized in that said monitoring method (30, 40) comprises the following steps: - detection (31, 41) of a regulation failure (PAN) affecting a regulation system (13, 14) of said at least one engine (3, 4), - detection (32, 42) of a current rotation speed (NR) of said at least one lift rotor (5), - determination (33, 43) of a compatibility condition (COMP) or an incompatibility condition (INCOMP) between a current flight phase and said rotation speed (NR), - in a first operating mode (M0D1), in the presence of said compatibility condition (COMP) and said regulation failure (PAN),generation (34, 44) of a first alert simultaneously representing said regulation failure (PAN) and said compatibility condition (COMP), and - in a second operating mode (M0D2), in the presence of said incompatibility condition (INCOMP) and said regulation failure (PAN), generation (35, 45) of a second alert simultaneously representing said regulation failure (PAN) and said incompatibility condition (INCOMP).,

2. Method according to claim 1, characterized in that said compatibility condition (COMP) is determined when said current rotation speed (NR) belongs to an interval of values defined by a rotation speed setpoint (NRcons) minus a first margin (n) and said rotation speed setpoint (NRcons) plus a second margin (m) and alternatively said incompatibility condition (INCOMP) is determined when said current rotational speed (NR) is excluded from said range of values.

3. Method according to claim 2, characterized in that said rotation speed setpoint (NRcons) is variable as a function of said current flight phase.

4. Method according to any one of claims 2 to 3, characterized in that said first margin (n) and said second margin (m) are variable depending on said current flight phase.

5. Method according to claim 1, characterized in that said compatibility condition (COMP) is determined when said current rotation speed (NR) belongs to a predetermined range of acceptable values allowing said current flight phase to continue and alternatively said incompatibility condition (INCOMP) is determined when said current rotation speed (NR) is excluded from said predetermined range of acceptable values allowing said current flight phase to continue.

6. Method according to claim 5, characterized in that said predetermined range comprises an upper limit defined so that a tangential speed at the tip of the blade of a blade (12) of said at least one lift rotor (5) is kept lower than the speed of sound.

7. A method according to any one of claims 5 to 6, characterized in that said predetermined range comprises a lower limit defined to provide a minimum thrust allowing said rotorcraft (1) to perform a flight at a constant altitude at a cruising forward speed.

8. Method according to any one of claims 1 to 7, characterized in that a transition from said first operating mode (M0D1) to said second operating mode (M0D2) is irreversible.

9. Method according to any one of claims 1 to 8, characterized in that said generation (34, 44) of the first alert comprises a first display (341, 441) on a display (10) of at least one item of information represented with a first predetermined color.

10. Method according to claim 9, characterized in that said generation (35, 45) of the second alert comprises a second display (351, 451) on said display (10) of said at least one piece of information represented with a second predetermined color different from said first predetermined color.

11. Method according to any one of claims 1 to 10, characterized in that, said at least one engine (3, 4) comprising a first engine (3) and a second engine (4), said regulation failure (PAN) affecting a first regulation system (13) of said first engine (3), when said second alert is generated, the method (30) comprises a command (36) for a shutdown of said first engine (3).

12. Method according to any one of claims 1 to 10, characterized in that, said at least one engine (3, 4) comprising a first engine (3) and a second engine (4), said regulation failure (PAN) affecting a first regulation system (13) of said first engine (3), when said second alert is generated, the method (40) comprises a control (46) of a reversible transmission device to prevent the transmission of engine torque from said first engine (3) to a power transmission chain (11).

13. Rotorcraft (1) comprising at least one lift rotor (5) driven in rotation by a power plant (2), said power plant (2) comprising at least one engine (3, 4), characterized in that said rotorcraft (1) comprises a monitoring system (6) comprising: - at least one failure sensor (23, 24) detecting a regulation failure (PAN) affecting a regulation system (13, 14) of said at least one engine (3, 4), - a speed sensor (7) measuring a current rotation speed (NR) of said at least one lift rotor (5), - a controller (8) determining a compatibility condition (COMP) or an incompatibility condition (INCOMP) between a current flight phase and said rotation speed (NR), said controller (8) generating, in a first operating mode (M0D1), in the presence of said compatibility condition (COMP) and said regulation failure (PAN),a first alert simultaneously representing said regulation failure (PAN) and said compatibility condition (COMP), and said monitoring controller (8) generating, in a second operating mode (MOD2), in the presence of said condition, incompatibility condition (INCOMP) and said regulation failure (PAN), a second alert simultaneously representing said regulation failure (PAN) and said incompatibility condition (INCOMP).

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