METHOD FOR MONITORING THE FUNCTION OF A TURNTABLE DRIVE SYSTEM AND ASSOCIATED TURNTABLE

Description

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

[0002] A rotorcraft has at least one rotor that contributes to its lift, or even its propulsion.

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

[0004] To rotate the rotor(s), the rotorcraft includes a propulsion system, 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 unit, which rotates the rotor(s). The power transmission unit then has one mechanical input per motor, with the mechanical inputs meshing with a combiner, the combiner driving one mechanical output of the power transmission unit per rotor. The motor inputs, the combiner, and the mechanical outputs can include at least one pinion or gear, at least one shaft, at least one speed reduction stage, etc.

[0006] Engines can be internal combustion engines with an output shaft set in motion by the combustion of fuel. For example, at least one engine can be a turboshaft engine equipped with a gas generator and a free-working turbine connected to the output shaft.

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

[0008] This engine control unit is then configured to control a fuel metering system supplying the associated engine in order, for example, to stabilize the output shaft's rotational speed at a set point, regardless of the power consumed by that output shaft. This power varies according to the 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 can be operated by an autopilot.

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

[0010] However, the engine control unit (ECU) is configured to prevent the power developed by an engine from exceeding a mechanical limit acceptable to the engine or the power transmission system. Typically, various operating modes, each with its own limits and durations, are then defined. For each operating mode, these limits might include, for example, a torque limit for the engine's output, an internal engine temperature limit, a rotational speed limit for an engine shaft, a torque limit for at least one component of the power transmission system, and so on.

[0011] On a multi-engine aircraft, there are two operating modes known as "AEO" (All Engines Operative), which can be used for certain periods when all engines are functioning normally, and two operating modes known as "OEI" (One Engine Inoperative), which can be used for certain periods when one of the engines is out of service. The use of certain operating modes may necessitate maintenance on the engines, or even on the power transmission system.

[0012] Therefore, an engine control unit (ECU) manages a fuel metering system to modulate the power of the associated engine. The fuel metering system is configured to bring the output shaft speed towards a set value, while preventing it from exceeding a limit.

[0013] Furthermore, in the event of a failure or malfunction of an engine's control system, the fuel flow to the affected engine is fixed at the last flow rate used. A pilot is notified and can then either shut down the affected engine 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] Documents GB2079707 and US 2011 / 173988 disclose engine control systems, such as those for a gas turbine, used in the event of a failure of another engine. Such systems, however, are unrelated to the invention.

[0016] The present invention aims to provide a method for limiting the workload of a crew in the event of a failure of an engine control system on a rotorcraft or in the absence of control of one of the engines. Furthermore, this method and the associated rotorcraft allow a pilot to perform piloting maneuvers safely, without risking damage to the mechanical power transmission system or a rotor of the rotorcraft.

[0017] The invention thus relates to a method for monitoring the operation of a rotorcraft propulsion system, the propulsion system comprising at least one engine, the rotorcraft comprising at least one lift rotor driven in rotation by the propulsion system, said monitoring method comprising the following steps: detection of a control failure affecting a control system of said at least one engine, detection of a current rotational 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 rotational speed NR.

[0018] According to the invention, such a monitoring method is remarkable in that it comprises the following steps: in a first mode of operation, in the presence of the compatibility condition and the regulation failure, generation of a first alert representing simultaneously the regulation failure and the compatibility condition, and in a second mode of operation, in the presence of the incompatibility condition and the regulation failure, generation of a second alert representing simultaneously the regulation failure and the incompatibility condition.

[0019] Furthermore, in the event of an engine control failure, the fuel metering unit supplying fuel to the associated engine's combustion chamber remains stuck in a current position and continues to deliver a constant fuel flow. Depending on the current flight phase, and particularly a change in rotor blade pitch control, the rotor's rotational speed (NR) can then evolve uncontrollably and fall outside its nominal operating range.

[0020] In addition, the current flight phase can 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.

[0021] In the first mode of operation, such a monitoring process allows a pilot to continue a phase of flight despite the failure of regulation of at least one engine, as long as the rotational speed NR is maintained within a nominal operating range for the current phase of flight.

[0022] The first alert is thus generated for informational purposes only, and the pilot can continue his mission without any particular action to take.

[0023] In the second operating mode, if the control failure of at least one engine persists and the RPM (reverse rotational speed) falls outside the nominal operating range for the current flight phase, then the second alert may allow the pilot to perform a predetermined emergency procedure. For example, the pilot may take an action to bring the RPM back within its nominal range or shut down the engine affected by the control failure.

[0024] Such a second alert is distinct from the first alert to allow a pilot to differentiate between them.

[0025] In addition, each alert can take the form of a visual alarm, for example by means of the emission of a light with a light-emitting diode or equivalent or the display on a screen of one or more characters, an audible alarm, by means of a loudspeaker, and / or a haptic alarm, for example by means of a vibrating unit causing an organ held or worn by an individual to vibrate.

[0026] According to a first embodiment of the invention, the compatibility condition can be determined when the current rotational speed NR belongs to an interval of values ​​defined by a rotational speed setpoint less a first margin (n) and the rotational speed setpoint plus a second margin (m) and alternatively the incompatibility condition can be determined when the current rotational speed NR is excluded from this interval of values.

[0027] In other words, in this case, the compatibility condition and the incompatibility condition are functions of a current rotational speed setpoint, a first margin (n) and a second margin (m). The first margin (n) and the second margin (m) can, for example, consist of a percentage of the rotational speed setpoint.

[0028] In practice, the rotation speed setting can vary depending on the current flight phase.

[0029] Such a rotation speed instruction can be generated by a flight computer of a rotorcraft's avionics system based in particular on atmospheric conditions of the flight phase and control instructions allowing the pilot to control the rotorcraft.

[0030] Similarly, the first margin (n) and the second margin (m) can vary depending on the current flight phase.

[0031] Indeed, depending on the current flight phase, the first and second margins (n) and (m) can be modified and thus allow the limits of the nominal operating range to be adapted to the acceptable values ​​for the rotor rotation speed NR.

[0032] According to a second 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.

[0033] In this case, the limits of the nominal operating range corresponding to the acceptable values ​​for the rotor rotation speed NR are fixed and predefined, for example, by ground tests, flight tests or simulations.

[0034] In this case, the predetermined range may include a defined upper bound to ensure that a tangential velocity at the blade tip of a blade of said at least one lift rotor is kept below the speed of sound.

[0035] Such an upper limit can thus ensure the integrity of the lift rotor. Alternatively or additionally, this upper limit can be defined to provide the rotorcraft with maximum performance during a hovering or low-speed flight phase and / or to limit the acoustic footprint generated by the rotor's rotation in the rotorcraft cabin.

[0036] Similarly, the predetermined range may include a defined lower bound to provide minimum thrust enabling the rotorcraft to fly at a constant altitude at a cruising forward speed.

[0037] 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 related to the rotor speed (NR) and increases as the rotor speed (NR) increases.

[0038] Alternatively or additionally, this lower limit can be defined to avoid an excessive drop in altitude of the rotorcraft in the event of an engine failure and before reaching a rotor speed that allows for an autorotation flight phase.

[0039] In practice, a switch from the first operating mode to the second operating mode can be irreversible.

[0040] Such irreversibility is indeed an additional safeguard. It contributes to flight safety and minimizes the risk of confusion for the pilot.

[0041] Advantageously, the generation of the first alert can include an initial display on a screen of at least one piece of information represented with a predetermined first color.

[0042] This 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.

[0043] The display may, for example, include a screen from the instrument panel of the rotorcraft or a so-called "head-up display" which can be worn by the pilot and integrated into glasses or a helmet screen.

[0044] The first predetermined color could be, for example, amber or orange.

[0045] Similarly, the generation of the second alert may include a second display on the display of said at least one piece of information represented with a second predetermined color different from the first predetermined color.

[0046] This 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.

[0047] The message and / or 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.

[0048] Furthermore, said at least one motor may, for example, comprise a first motor and a second motor. In this case, during a switchover to the second operating mode, the method may implement additional steps aimed at bringing the rotor's rotational speed NR back into its nominal operating range.

[0049] Thus, according to a first variant, the regulation failure affecting a first regulation system of the first motor, when the second alert is generated, the process may include a command to stop the first motor.

[0050] Such a command to stop the first engine could, for example, consist of closing a fuel supply solenoid valve on that engine. Once the solenoid valve is closed, a first fuel metering unit connected to the combustion chamber of the first engine can no longer supply fuel, and the first engine stops. However, the rotor continues to rotate thanks to the second engine, whose control system is not affected by the failure.

[0051] According to a second variant, the regulation failure affecting a first regulation system of the first motor, when the second alert is generated, the method may include a control of a reversible transmission device to prevent the transmission of motor torque from the first motor to a power transmission chain.

[0052] Such a control of a reversible transmission device for the first engine could, for example, consist of disengaging a clutch system or a piloted freewheel arranged between an output shaft of the first engine and an input shaft of a power transmission gearbox. In this case, the first engine does not stop, but it no longer supplies engine torque to the rotor via the power transmission chain.

[0053] The present invention also relates to a rotorcraft comprising at least one lifting rotor driven in rotation by a drive system, the drive system comprising at least one engine, said rotorcraft comprising a monitoring system comprising: at least one fault sensor detecting a control failure affecting a control system of said at least one motor, a speed sensor measuring a current rotational 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 rotational speed NR.

[0054] According to the invention, such a rotorcraft is remarkable in that said controller generates, in a first mode of operation, in the presence of the compatibility condition and the regulation failure, a first alert representing simultaneously the regulation failure and the compatibility condition, and in that said controller generates, in a second mode of operation, in the presence of the incompatibility condition and the regulation failure, a second alert representing simultaneously the regulation failure and the incompatibility condition.

[0055] Furthermore, such a fault sensor can be integrated into the control system of at least one motor, which is connected to the controller via wired or wireless means. The fault sensor thus transmits information about a fault in the relevant control system directly to the controller.

[0056] Similarly, the airspeed sensor can be integrated into an avionics system that is connected to the controller via wired or wireless means. Such an avionics system also transmits information representative of flight conditions to deduce a current flight phase.

[0057] From the information received, the controller deduces the compatibility condition or the incompatibility condition between the current flight phase and the rotation speed NR.

[0058] Upon receiving information about the regulation failure, the controller can then determine whether to implement the first operating mode or the second operating mode and thus generate the corresponding alert.

[0059] The invention and its advantages will become apparent in more detail in the following description, with illustrative examples given by reference to the attached figures which represent: therefigure 1 , a diagram of a rotorcraft in side view equipped with a monitoring system according to the invention, the figure 2 , a flowchart illustrating a monitoring method according to the invention, the figure 3 , a flowchart illustrating a first alternative to the monitoring method according to the invention, and the figure 4 , a flowchart illustrating a second alternative of the monitoring method according to the invention.

[0060] Elements present in several separate figures are assigned a single reference.

[0061] As already mentioned, the invention relates to a rotorcraft equipped with a monitoring system for a propulsion installation and the associated monitoring method.

[0062] As depicted in the figure 1 , such a rotorcraft 1 comprises a propulsion system 2 equipped with at least one motor 3, 4 enabling 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 motor 3, 4.

[0063] Such a power transmission box 9 then includes a mechanical output intended to drive in rotation a rotor mast fixed in rotation to said at least one lift rotor 5.

[0064] In addition, the fuel supply flow 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.

[0065] The monitoring system 6 then includes at least one fault sensor 23, 24 capable of detecting a PAN regulation fault affecting the regulation system 13, 14 of said at least one motor 3, 4.

[0066] The monitoring system 6 also includes a speed sensor 7, for example arranged at the rotor mast, and capable of measuring a current rotational speed NR of said at least one lift rotor 5.

[0067] The terms "fault sensor 23, 24" and "speed sensor 7" refer to physical sensors capable of directly measuring the parameter in question, as well as a system that may include one or more physical sensors and signal processing means for providing an estimate of the parameter based on the measurements provided by these physical sensors. Similarly, the term "measurement of this parameter" will refer to both a raw measurement from a physical sensor and a measurement obtained through more or less complex signal processing from one or more raw measurements.

[0068] The monitoring system 6 then includes a controller 8 connected by wired or wireless means with the speed sensor 7 and a set of other sensors not shown, for example from an avionics system of the rotorcraft 1, configured to allow identification of a current flight phase of the rotorcraft 1.

[0069] The set of sensors configured to identify a current flight phase may include, but is not limited to, an aerodynamic data computer (ADC), an attitude and heading reference system (AHRS), a Rad Alt radar altimeter, and a GPS global satellite positioning system.

[0070] The ADC aerodynamic data computer provides the monitoring system 6 in the non-limiting embodiment described, but without limitation, with the barometric altitude, the speed of movement of the rotorcraft 1 relative to the air and the gross vertical speed of the rotorcraft 1.

[0071] The AHRS heading and attitude reference system provides the monitoring system 6 in the non-limiting embodiment described, but not limited to, attitude and heading, acceleration and rate of descent information.

[0072] With input from the ADC aerodynamic data computer, the AHRS heading and attitude reference system provides an instantaneous vertical speed.

[0073] With input from a GPS global positioning satellite system and / or an FMS flight avionics system, the AHRS heading and attitude reference system provides a combined navigation position and ground speed in each direction.

[0074] The Rad Alt radar altimeter provides the height of the rotorcraft 1 above ground and water.

[0075] Controller 8 thus allows determining a COMP compatibility condition or an INCOMP incompatibility condition between a current flight phase and the rotation speed NR.

[0076] Furthermore, controller 8 may include, 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 do not limit the scope given to the term "controller." The term "processor" may refer to a central processing unit known by the acronym CPU, a graphics processing unit (GPU), a digital signal processing unit (DSP), a microcontroller, etc.

[0077] In addition, the monitoring system 6 may also include a display 10 enabling the display to a pilot of the rotorcraft 1 of two separate warning messages when a PAN regulation failure is detected.

[0078] Furthermore, according to the figures 2 à 4 , the invention also relates to a method of monitoring 30, 40 the operation of the drive installation 2 which drives in rotation the lift rotor(s) 5.

[0079] Furthermore, instructions or a computer program can be stored in a memory of the monitoring system 6. The monitoring system 6 can then execute these instructions or this program to implement the monitoring process 30, 40.

[0080] Such a monitoring method 30, 40 thus includes a detection 31, 41 of a PAN regulation failure affecting a regulation system 13, 14 of said at least one motor 3, 4.

[0081] The monitoring method 30, 40 also includes detecting 32, 42 the current rotational speed NR of at least one lift rotor 5 and determining 33, 43 a compatibility condition COMP or an incompatibility condition INCOMP between a current flight phase and the rotational speed NR. Consequently, the compatibility condition COMP and the incompatibility condition INCOMP are two opposing conditions that cannot be verified simultaneously.

[0082] Therefore, in a first operating mode MOD1, in the presence of the COMP compatibility condition and the PAN regulation failure, the monitoring process 30, 40 includes a generation 34, 44 of a first alert representing simultaneously the PAN regulation failure and the COMP compatibility condition.

[0083] Alternatively, and in a second operating mode MOD2, in the presence of the incompatibility condition INCOMP and the PAN regulation failure, the monitoring process 30, 40 includes a generation 35, 45 of a second alert representing simultaneously the PAN regulation failure and the incompatibility condition INCOMP.

[0084] Furthermore, the transition from the first operating mode MOD1 to the second operating mode MOD2 may be irreversible for flight safety reasons.

[0085] The generation steps 34, 44 of a first alert and of generation 35, 45 of a second alert are thus implemented by the controller 8 who is able to generate at least two alerts distinct from each other and thus inform the pilot of the gyroplane 1 of a current safety level allowing or prohibiting the continuation of a mission during a PAN regulation failure.

[0086] In addition, various alternatives allow for determining the conditions of COMP compatibility and INCOMP incompatibility.

[0087] According to a first alternative of the monitoring process 30 represented at the figure 3 , the compatibility condition COMP is determined for example when the current rotational speed NR belongs to an interval of values ​​defined by a rotational speed setpoint NRcons less a first margin n and the rotational speed setpoint NRcons plus a second margin m and alternatively the incompatibility condition INCOMP is determined when the current rotational speed NR is excluded from this interval of values.

[0088] In practice, such a rotation speed setpoint NRcons can be variable depending on the current flight phase identified by the controller 8.

[0089] Similarly, the first margin n and the second margin m can also vary depending on the current flight phase.

[0090] Furthermore, the generation 34 of the first alert may include a first display 341 on the display 10 of at least one piece of information represented with a first predetermined colour.

[0091] The generation 35 of the second alert may, for its part, include a second display 351 on the display 10 of said at least one piece of information represented with a second predetermined color different from the first predetermined color.

[0092] For example, the first predetermined color could be amber and the second predetermined color could be red.

[0093] The said information may for example include an alphanumeric message “FAIL” indicating a PAN regulation failure and being completed with the relevant regulation system 13 or 14 “FADEC1” or “FADEC2”.

[0094] Furthermore, when the rotorcraft 1 has 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 process 30 may include a command 36 to stop the first engine 3.

[0095] 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.

[0096] According to a second alternative of the monitoring process 40 represented in the figure 4 The COMP compatibility condition can be determined when the current NR rotation speed belongs to a predetermined range of acceptable values ​​allowing the current flight phase to continue, and alternatively the INCOMP incompatibility condition can be determined when the current NR rotation speed is excluded from this predetermined range of acceptable values.

[0097] For example, the predetermined range may include a defined upper bound so that a tangential velocity at the blade tip of a blade 12 of said at least one lift rotor 5 is kept below the speed of sound.

[0098] Similarly, the predetermined range may include a lower bound defined to provide a minimum thrust enabling the rotorcraft 1 to fly at a constant altitude at a cruising forward speed.

[0099] Furthermore, the generation 44 of the first alert may include a first display 441 on the display 10 of at least one piece of information represented with a first predetermined colour.

[0100] The second alert generation 45 may, for its part, include a second display 451 on the display 10 of said at least one piece of information represented with a second predetermined color different from the first predetermined color.

[0101] The said 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 band in red to represent the second alert.

[0102] In addition, such a luminous geometric shape can be displayed as a background of a digital indicator showing a current value of the rotational speed NR.

[0103] Furthermore, when the rotorcraft 1 has 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 include a control 46 of a reversible transmission device to prevent the transmission of an engine torque from the first engine 3 to the power transmission chain 11.

[0104] 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.

[0105] Naturally, the present invention is subject to numerous variations in its implementation. Although several embodiments have been described, it is understood that it is not possible to exhaustively identify all possible embodiments. The scope of the present invention is defined by the statement of claims.

Claims

1. Method (30, 40) for monitoring an operation of a rotorcraft (1) engine installation (2), said engine installation (2) comprising at least one engine (3, 4), said rotorcraft (1) comprising at least one lift rotor (5) rotated by said engine installation (2), said monitoring method (30, 40) comprising the following steps: - detection (31, 41) of a regulation breakdown (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 of an incompatibility condition (INCOMP) between a current flight phase and said rotation speed (NR), characterised in that said monitoring method (30, 40) comprises the following steps: - in a first operating mode (MOD1), in the presence of said compatibility condition (COMP) and of said regulation breakdown (PAN), generation (34, 44) of a first alert, simultaneously representative of said regulation breakdown (PAN) and of said compatibility condition (COMP), and - in a second operating mode (MOD2), in the presence of said incompatibility condition (INCOMP) and of said regulation breakdown (PAN), generation (35, 45) of a second alert, simultaneously representative of said regulation breakdown (PAN) and of said incompatibility condition (INCOMP).

2. Method according to claim 1, characterised in that said compatibility condition (COMP) is determined when said current rotation speed (NR) belongs to a value interval defined by a rotation speed setpoint (NRcons) less 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 rotation speed (NR) is excluded from said value interval.

3. Method according to claim 2, characterised 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, characterised in that said first margin (n) and said second margin (m) are variable as a function of said current flight phase.

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

6. Method according to claim 5, characterised in that said predetermined range comprises a defined upper limit, such that a vane end tangential speed of a vane (12) of said at least one lift rotor (5) is maintained less than the speed of sound.

7. Method according to any one of claims 5 to 6, characterised in that said predetermined range comprises a defined lower limit to provide a minimal thrust, making it possible for said rotorcraft (1) to fly at a constant altitude, at a forward cruise speed.

8. Method according to any one of claims 1 to 7, characterised in that a passage from said first operating mode (MOD1) to said operating mode (MOD2) is irreversible.

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

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

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

12. Method according to any one of claims 1 to 10, characterised in that, said at least one engine (3, 4) comprising a first engine (3) and a second engine (4), said regulation breakdown (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 an engine torque of said first engine (3) to a power transmission chain (11).

13. Rotorcraft (1) comprising at least one lift rotor (5), rotated by an engine installation (2), said engine installation (2) comprising at least one engine (3, 4) said rotorcraft (1) comprising a monitoring system (6) comprising: - at least one breakdown sensor (23, 24) detecting a regulation breakdown (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), characterised in that said controller (8) generates, in a first operating mode (MOD1), in the presence of said compatibility condition (COMP) and of said regulation breakdown (PAN), a first alert, simultaneously representative of said regulation breakdown (PAN) and of said compatibility condition (COMP), and and in that said monitoring controller (8) generates, in a second operating mode (MOD2), in the presence of said incompatibility condition (INCOMP) and of said regulation breakdown (PAN), a second alert, simultaneously representative of said regulation breakdown (PAN) and of said incompatibility condition (INCOMP).