Method for reactivating internal combustion engine in standby mode during asymmetric mode of operation in multi-engine aircraft

Abstract

The invention relates to a method of driving a multi-engine aircraft (1) having an asymmetric operating mode, in which at least one engine (10) is an active engine regulated by a regulation system (55) by outputting non-zero active driving power participating in rotation of a rotor (5), and at least one engine (10) is a passive engine in a standby mode. During the asymmetric mode of operation, the method comprises the steps of: detecting, using a conditioning system (55), operation of the power plant (2) in a risky flight phase that is disadvantageous to exit from the asymmetric mode of operation prior to the flight phase; and reactivating the passive engine after detecting operation of the power plant (2) in the risk flight phase.

Description

Method for reactivating an internal combustion engine in standby mode during asymmetric operation in a multi-engine aircraft

[0001] Cross-references to related applications

[0002] This application claims the benefit of French patent application FR 24 11828, filed on October 29, 2024, the disclosure of which is incorporated herein by reference in its entirety. Technical Field

[0003] The present invention relates to a method for reactivating an internal combustion engine in standby mode during asymmetric operation in a multi-engine aircraft.

[0004] An aircraft may include multiple internal combustion engines for moving mechanical systems, such as those that rotate at least one rotor on a helicopter.

[0005] Internal combustion engines can take the form of turboshaft engines, which may have free turbines. A free turbine turboshaft engine includes a gas generator equipped with a compressor, a combustion chamber, and a high-pressure expander assembly constrained to rotate with the compressor. The compressor may have one or more compression stages. Similarly, the expander assembly may include one or more expander turbines. Furthermore, the free turbine turboshaft engine includes at least one low-pressure working turbine, i.e., a low-pressure working turbine that is mechanically independent of the compressor and the high-pressure expander assembly in rotation. The working turbine then rotates a power shaft connected to a mechanical system to be set to motion.

[0006] According to another example, an internal combustion engine can include a piston engine.

[0007] Therefore, a rotorcraft may include a power plant with multiple internal combustion engines for moving the mechanical system, and in particular a power transmission system for rotating at least one rotor.

[0008] Alternatively, the power unit can operate in a cooperative mode, in which each internal combustion engine generates a driving force that collectively propels the mechanical system into motion.

[0009] The power plant can also operate in an asymmetric mode by placing one of the engines in standby mode during certain operational phases of the aircraft. In rotorcraft, during asymmetric mode, at least one active engine is modulated to ensure rotor rotation by transmitting non-zero active drive power to the rotor via the rotor's power shaft. However, at least one passive engine can be in standby mode.

[0010] An engine in standby mode can be shut down by closing the combustion chamber. Then, the passive engine's combustion chamber is no longer supplied with fuel, and on the turboshaft engine, the rotating portion of the gas generator can be programmed to move by an electric motor or not. Alternatively, the engine in standby mode can operate in idle or over-idle mode, with the combustion chamber ignited and supplied with fuel. The engine in standby mode operates at very low speeds and does not transmit power to the rotor, while one or more active engines provide the entire power supply. For example, the passive engine in idle mode rotates at approximately 40% of its rated speed. Background Technology

[0011] Document EP 3209563 specifically describes the asymmetric operation mode.

[0012] To exit asymmetric operating mode, the passive engine must be reactivated by being controlled to increase power output. However, reactivation can be difficult under certain conditions.

[0013] Reference FR 3135965 proposes that when an internal combustion engine in standby mode is subjected to icing conditions, hot air from the active internal combustion engine can be used to heat the internal combustion engine in standby mode. Summary of the Invention

[0014] Therefore, the object of the present invention is to provide a method and an aircraft for optimizing the reactivation of an internal combustion engine in standby mode during asymmetric operation by following strategies other than those described in FR 3135965.

[0015] Therefore, the present invention relates to a method for piloting an aircraft having a power unit (2) comprising at least two internal combustion engines and a transmission system connected to at least one rotor, each internal combustion engine having a power shaft connected to the transmission system, the method comprising an asymmetric operating mode comprising using an adjustment system to actively adjust at least one active engine of the at least two internal combustion engines at an active speed, the active engine outputting non-zero active drive power at an active speed and via its power shaft to cause the rotor to rotate, the asymmetric operating mode comprising, together with adjusting the active engine at an active speed, placing at least one passive engine of the at least two internal combustion engines in a standby mode, the passive engine not transmitting power to the rotor.

[0016] Therefore, in the asymmetric operation mode, the method includes the following steps:

[0017] • Use the adjustment system to detect the operation of the power unit during a risky flight phase prior to the flight phase, an operation that is detrimental to exiting asymmetric operation mode and therefore detrimental to reactivating the passive engines to return to cooperative operation mode; and

[0018] • After detecting operation of the power unit during a risky flight phase, the passive engine is reactivated using the adjustment system, and then the passive engine becomes an active engine that adjusts its speed actively.

[0019] The control system may include part of the aircraft's avionics system.

[0020] Therefore, this method addresses the bias arising from document FR 3135965, which recommends preheating the engine in standby mode. Documents US2020256265A1, FR3135965A1, and WO2024161091A1 are also known.

[0021] In asymmetric operation, at least one internal combustion engine is intentionally placed in standby mode, with one or more active engines providing all power. The passive engine may have a closed combustion chamber, and on a turboshaft engine, the rotating components of the gas generator can be set to motion by an electric motor. Alternatively, the passive engine may have an ignited combustion chamber without transmitting drive power to the rotor. In this case, the passive engine has a low rotational speed compared to its rated speed, for example, approximately 40% of the rated speed.

[0022] Passive engines are ventilated, in which outside air continues to flow within the engine, especially when the aircraft includes dynamic air intakes.

[0023] Therefore, this invention proposes monitoring current flight conditions to determine whether the aircraft is operating during a risky flight phase, potentially leading to short-term operation under conditions unfavorable to exiting asymmetric operating mode. If so, the passive engine is reactivated before the unfavorable flight conditions are reached. If necessary, the combustion chamber is reignited. If the passive engine is in idling mode, it is reactivated by acceleration before the unfavorable flight conditions for reactivation are reached. Thus, this invention enables the safeguarding of asymmetric operating mode.

[0024] The method may also include one or more of the following features, used individually or in combination.

[0025] According to the first variant, after detecting operation of the power unit in a risky flight phase, the method may include issuing an alarm using an alarm device, and after manipulating the human-machine interface, the regulation system performs the reactivation of the passive engine.

[0026] In this scenario, the method warns the pilot that the current flight conditions necessitate exiting asymmetric mode. The pilot can then manipulate the human-machine interface for this purpose. This interface can also be used to activate asymmetric mode.

[0027] According to the second variant, the reactivation of the passive engine can be automatically performed by the control system after the operation of the power unit is detected in a risky flight phase.

[0028] In this situation, the control system automatically reactivates the passive engine. Alarms can be issued in parallel to warn the pilot to exit asymmetric operating mode.

[0029] Optionally, the aircraft may include a parameter setting interface that allows the pilot to choose whether to apply the first variant or the second variant.

[0030] According to one possibility compatible with the foregoing possibilities, the method may include using various sensors of the control system to measure at least one monitoring parameter, and detecting operation of the power unit in a risky flight phase may include detecting that the monitoring parameter has a current value that is less than an associated limit.

[0031] At least one aircraft monitoring parameter is measured and compared with its own limits to determine whether the aircraft should exit the asymmetric operating mode. For each monitoring parameter, associated limits can be established through testing, calculation, and / or simulation.

[0032] For example, the at least one monitoring parameter may include at least one of the following parameters:

[0033] ● Temperature value, which is a function of the external temperature of the air around the aircraft as measured using an external temperature sensor, and the associated limit is a stored external temperature limit. The operation of detecting a power unit in a risky flight phase includes detecting a temperature value that is less than the stored external temperature limit.

[0034] • The oil temperature of the lubrication circuit of the passive engine is measured using an oil temperature sensor, and the associated limit is a stored oil temperature limit. Detecting operation of the power unit during a risky flight phase includes detecting that the oil temperature has a current value lower than the stored oil temperature limit; and

[0035] ● The fuel temperature of the fuel supplied to the passive engine is measured using a fuel temperature sensor, and the associated limit is a stored fuel temperature limit. The operation of detecting a power unit in a risky flight phase includes detecting that the fuel temperature has a current value that is lower than the stored fuel temperature limit.

[0036] Optionally, all three referenced parameters can be monitored simultaneously.

[0037] In reality, outside air circulates within the passive engine. Therefore, the temperature of this outside air affects the temperature of the passive engine components.

[0038] Environmental conditions also affect engine oil and fuel temperatures. When oil and / or fuel temperatures are particularly low, restarting a passive engine can be difficult.

[0039] Alternatively, the temperature value mentioned above can be equal to the external temperature measured using an external temperature sensor.

[0040] Alternatively, the aforementioned temperature value can be calculated by the control system based on the external temperature and the aircraft's current speed as measured by a speed sensor.

[0041] For example, such current speed could be the airspeed of the aircraft. The higher the current speed of the aircraft, the more fresh outside air is supplied to the passive engine. Therefore, the regulation system could include mathematical laws that provide temperature values ​​based not only on the outside temperature but also on the current speed, in order to account for the possibility that this air supply might cool the passive engine.

[0042] As a supplement or alternative, the at least one monitoring parameter may include the ambient temperature in the engine compartment housing the passive engine, measured using an ambient temperature sensor, and the associated limit is a stored ambient temperature limit. Detecting the operation of the power unit in a risky flight phase includes detecting that the ambient temperature has a current value that is less than the stored ambient temperature limit.

[0043] In fact, the prevailing temperature in the engine compartment of a passive engine can affect the engine's operation.

[0044] As a supplement or alternative, the at least one monitoring parameter may include the internal temperature in the passive engine measured using an internal temperature sensor, and the associated limit is a stored internal temperature limit. Detecting the power unit in a risky flight phase includes detecting that the internal temperature has a current value that is less than the stored internal temperature limit.

[0045] For example, the temperature in the combustion chamber can be measured, and reactivation may be difficult below associated limits.

[0046] In a supplementary or alternative manner, or even specifically, if the passive engine is in idling mode, the at least one monitoring parameter may include the oil pressure of the lubrication circuit of the passive engine, which is measured using an oil pressure sensor, and the associated limit is a stored oil pressure limit. Detecting operation of the power unit in a risky flight phase includes detecting that the oil pressure has a current value that is less than the stored oil pressure limit.

[0047] In a supplementary or alternative manner, or even specifically, if the passive engine is in idling mode, the at least one monitoring parameter may include the fuel pressure of the fuel supplied to the passive engine, and the associated limit is a stored fuel pressure limit, which is measured using a fuel pressure sensor. Detecting operation of the power unit in a risky flight phase includes detecting that the fuel pressure has a current value that is less than the stored fuel pressure limit.

[0048] According to one possibility compatible with the foregoing possibilities, the driving method according to the invention may include: issuing an alarm using an alarm device when the at least one monitored parameter is greater than or equal to the associated limit and less than or equal to the associated threshold, the alarm signaling the aircraft's operation during a transient phase, wherein the associated threshold is greater than the associated limit.

[0049] A second limit can be implemented for each monitored parameter. This second limit, referred to as a “threshold,” is used to distinguish it from the limit that necessarily means exiting asymmetric operating mode. An alarm is issued when a monitored parameter has a value between the associated threshold and limit. This alarm can be used to signal to the pilot that the aircraft is approaching flight conditions that require exiting asymmetric operating mode. The pilot can then choose to exit these flight conditions, for example, by slowing down and / or decreasing the aircraft’s altitude, or simply by paying closer attention to the changing flight conditions.

[0050] The present invention also relates to an aircraft having a power plant comprising at least two internal combustion engines and a transmission system connected to at least one rotor, each internal combustion engine having a power shaft connected to the transmission system, the aircraft including an asymmetric operating mode comprising using an adjustment system to actively adjust at least one active engine of the at least two internal combustion engines at an active speed, the active engine outputting non-zero active drive power at an active speed and via its power shaft to cause the rotor to rotate, the asymmetric operating mode comprising, together with adjusting the active engine at an active speed, placing at least one passive engine of the at least two internal combustion engines in a standby mode, the passive engine not transmitting power to the rotor.

[0051] The adjustment system is then configured to implement the method of the present invention.

[0052] For this purpose, the control system may include at least one controller configured to: i) detect operation of the power unit in a risky flight phase, for example by comparing the current value of one or more monitoring parameters with one or more respective limits; and ii) reactivate the passive engine and thus exit asymmetric mode after detecting operation of the power unit in a risky flight phase.

[0053] For this purpose, the control system may include at least one of the following sensors: an external temperature sensor that measures the external temperature of the air surrounding the aircraft; an oil temperature sensor for each internal combustion engine that measures the oil temperature circulating in the internal combustion engine; a fuel temperature sensor (which may be universal or one for each engine) that measures the fuel temperature supplied to each internal combustion engine; a speed sensor that measures the current speed of the aircraft; an ambient temperature sensor for each internal combustion engine that measures the ambient temperature in the compartment housing the internal combustion engines; an internal temperature sensor for each internal combustion engine that measures the internal temperature of the internal combustion engines; an oil pressure sensor for each internal combustion engine that measures the internal temperature of the internal combustion engines; and a fuel pressure sensor for each internal combustion engine that measures the fuel pressure of the fuel circulating in the internal combustion engines.

[0054] The control system may also include an alarm. Attached Figure Description

[0055] The invention and its advantages will become more apparent from the following description of examples given by way of illustration with reference to the accompanying drawings, wherein:

[0056] Figure 1 is a view of the aircraft according to the present invention; and

[0057] Figure 2 is a diagram illustrating the method used. Detailed Implementation

[0058] Elements present in more than one figure are given the same reference numerals in each of them.

[0059] Figure 1 shows an example of an aircraft 1 according to the present invention. The aircraft 1 includes a rotor 5. The rotor 5 is equipped with a plurality of blades 6, which are rotatable and can be supported by a hub 7 or equivalent. For example, the rotor 5 forms a fixed or tilting propeller, rotor, or anti-torque rotor.

[0060] The aircraft 1 has a power unit 2 for rotating the rotor 5. The power unit 2 is equipped with at least two internal combustion engines 10. Reference numeral 10 indicates any internal combustion engine, and reference numerals 11 and 12 may indicate specific engines if necessary. Each internal combustion engine 10 is housed in an engine nacelle 100. Reference numeral 100 indicates any engine nacelle, and reference numerals 101 and 102 indicate the engine nacelles of internal combustion engines 11 and 12, respectively.

[0061] According to one example, at least one internal combustion engine 10 may be a turboshaft engine. Such a turboshaft engine 10 includes a gas generator 15 equipped with at least one compression turbine 16, a combustion chamber 17 into which fuel is injected, and at least one expansion turbine 18 constrained to rotate together with one or more compression turbines 16. In addition, the turboshaft engine may include at least one operating turbine 19 that directly or indirectly sets the power shaft 20 of the engine 10 to motion.

[0062] Alternatively, at least one internal combustion engine 10 may be a piston engine equipped with a combustion chamber and a power shaft.

[0063] Regardless of the type of engine, each internal combustion engine 10 therefore includes a power shaft 20 connected to a power transmission system 25. The power transmission system 25 is then connected to the rotor 5 in the usual manner. Reference numeral 20 indicates any power shaft, and reference numerals 21 and 22 indicate specific power shafts of the two engines 11 and 12, respectively.

[0064] As an illustration, the power transmission system 25 may be equipped with a power transmission gearbox 26, which is mechanically inserted between the engine 10 and the rotor 5. The power transmission gearbox 26 may be configured such that each engine 10 is equipped with an input shaft 30, and various gears are arranged between the input shaft 30 and the rotor mast 35 fixed to the hub 7.

[0065] The power transmission system 25 may include, for example, at least one flywheel 50, 51, 52 between each engine 10 and the power transmission gearbox 26, and / or at least one connecting shaft 53, 54, and / or at least one connector that allows for misalignment. Various types of power transmission gearboxes and various mechanical systems are described in the literature, and the examples described are given by way of illustration only.

[0066] Furthermore, engine 10 is a fuel-powered internal combustion engine and can be started by a starter motor (not shown for clarity). Therefore, aircraft 1 includes a regulation system 55 for guiding the starter motor and the power output of each engine 10 via its power shaft 20.

[0067] Therefore, the regulating system 55 includes a fuel metering device 69 for each engine 10. Each engine 10 is then connected to at least one fuel tank 70 via its own fuel metering device 69. Reference numeral 69 indicates any fuel metering device, and reference numerals 71 and 72 indicate specific fuel metering devices for the two engines 11 and 12, respectively.

[0068] The control system 55 may include an engine computer 60 for each engine 10. Each engine computer 60 may include a processing unit. Such a processing unit may have, for example, at least one processor 64 and at least one memory 65, at least one integrated circuit, at least one programmable system, or at least one logic circuit; these examples do not limit the scope to the term "processing unit". The engine computers 60 may communicate with each other via wired or wireless links.

[0069] According to the described example, the power unit 2 includes two engine computers 61 and 62 that control two engines 11 and 12 respectively. Each engine computer 60 is configured to direct the associated engine 10 and operate it at a desired speed. In particular, each engine computer 61 and 62 can control the fuel metering devices 71 and 72 of the engine 11 and 12. Each engine computer 61 and 62 can be connected to multiple sensors to control the associated engine 11 and 12, such as, for example, a temperature sensor 930 that measures, for example, the gas temperature at the inlet of a free turbine, a speed sensor that measures, for example, the rotational speed of the gas generator of a turbine shaft engine, torque meters 910 and 920 that measure the engine torque on the rotating component, speed sensors 940 and 950 that measure, for example, the rotational speed of the rotating component, a torque meter that measures the torque applied to the rotor mast 35, a speed sensor that measures, for example, the rotational speed of the rotor mast 35, a sensor that measures external pressure, a sensor that measures external temperature, etc.

[0070] The engine computer 60 can form a controller 75 for applying the methods of the present invention, or the management computer 78 of the regulation system 55 can act as the controller 75. For example, the management computer 78 may include at least one processing unit. The management computer 78 can communicate with each engine computer 60 via a wired or wireless link. The management computer 78 can communicate directly or via the engine computers with each of the aforementioned measurement systems.

[0071] Regardless of its composition, the controller 75 can communicate with at least one siren 80 via a wired or wireless link to provide information to the pilot. For example, the siren 80 may include a display capable of showing messages, light-emitting diodes that illuminate upon commands from the controller 75, and a speaker.

[0072] In addition, the controller 75 can communicate with the human-machine interface 81.

[0073] Additionally, according to the given example, the controller 75 can communicate with at least one human-machine interface 85, or even with one human-machine interface for each engine 10 (i.e., two interfaces 86, 87). Each human-machine interface 85 can, for example, issue signals carrying stop commands, idle commands, or flight speed commands for the relevant engine 10. Figure 1 illustrates the interfaces in the three-position stop POS1 / idle POS2 / flight POS3 for this purpose.

[0074] In addition, the controller 75 can communicate with the human-machine interface 88, which can be manipulated to manually activate or deactivate the asymmetric operation mode.

[0075] Each human-machine interface 81, 85, 88 may include equipment that can be operated by the pilot, such as buttons or levers, touchscreens, voice commands, etc. According to the example shown, human-machine interfaces 81, 85, 88 and alarm 80 communicate with management computer 78. Alternatively or additionally, human-machine interfaces 81, 85, 88 and alarm 80 communicate with one or each engine computer unit 60.

[0076] Furthermore, the controller 75 can communicate directly or via the engine computer with multiple sensors 40 to 47, 410, 440, 450, 460, and 470. The term "sensor" should be understood to mean a physical sensor capable of directly measuring the parameter in question, and also to mean a system that may include one or more physical sensors, and means for processing the signals so that estimates of the parameter can be provided based on the measurements provided by these physical sensors. Similarly, the concept of a measured parameter refers to both the raw measurement from the physical sensors and the measurement obtained through relatively complex processing of the raw measurement signals.

[0077] Therefore, controller 75 can communicate with one or more of the following sensors:

[0078] ● External temperature sensor 40 measures the external temperature T0 of the air surrounding the aircraft 1;

[0079] • Each internal combustion engine 10 has its own oil temperature sensor 41, 410 to measure the oil temperature TOIL in the lubrication circuit of the relevant internal combustion engine 10;

[0080] • At least one fuel temperature sensor 42 for measuring the fuel temperature of the TFUEL, the system being capable of having a single sensor for measuring the temperature of the fuel in the fuel tank 70 or a separate sensor for measuring the fuel temperature of each engine.

[0081] • A speed sensor 43 that measures the current speed of the aircraft 1; such a speed sensor may include, for example, a satellite positioning system and / or a pitot tube system;

[0082] • Each engine 10 has its own ambient temperature sensor 44, 440, which measures the ambient temperature TCOMP in the engine compartment 100;

[0083] • Each internal combustion engine 10 has its own internal temperature sensor 45, 450, which measures the internal temperature LTENG in the internal combustion engine 10, such as the temperature of “T45” upstream of the free turbine in a turboshaft engine, as referred to by those skilled in the art.

[0084] • Each internal combustion engine 10 has its own oil pressure sensor 46, 460, which measures the oil pressure circulating in the lubrication circuit of the internal combustion engine 10; and

[0085] Each internal combustion engine 10 has its own fuel pressure sensor 47, which measures the fuel pressure of the fuel circulating in the internal combustion engine 10.

[0086] Figure 2 illustrates a driving method according to the present invention, which can be implemented by a rotorcraft 1 of the type shown in Figure 1. The method is illustrated using the control system 55 of Figure 1. However, the method is applicable to control systems without a management computer 78, as the engine computer 60 can be easily configured to apply the method.

[0087] During cooperative operation, each internal combustion engine 10 is a so-called active engine, regulated by the control system 55 to output non-zero active drive power via its power shaft 20 to cause the rotor 5 to rotate. For example, a first human-machine interface 86 is positioned at POS3 and transmits control signals to the engine computer 61. Similarly, a second human-machine interface 87 is positioned at POS3 and transmits control signals to the engine computer 62.

[0088] To activate the asymmetric operating mode, the pilot can manipulate the human-machine interface 88. The control system keeps at least one of the engines, referred to as the "active engine," active and at least one of the engines, referred to as the "passive engine," in standby mode. One or more passive engines then shut down, having closed combustion chambers, or idle so that power is not transmitted via the associated flywheel.

[0089] Optionally, the method may include the step of using the adjustment system 55 to detect that the STPASY asymmetric operation mode is activated, for example by detecting a signal carrying a command to apply the asymmetric operation mode issued by the human-machine interface 88.

[0090] When the asymmetric operating mode is activated, the method may include using the regulation system 55 to detect the operation of the power unit 2 in a risky flight phase that is unfavorable to the reactivation of the passive engine before the STPD flight phase.

[0091] In some embodiments, the detection STPD of the power unit 2 during a risky flight phase, prior to the flight phase, which is unfavorable for exiting the asymmetric operating mode, is operated using the regulation system 55 independently of the operation of the at least one active engine among the at least two internal combustion engines 10.

[0092] Therefore, the method may involve using one or more of the various sensors 40 to 47, 410, 440, 450, 460, 470 to measure one or more monitoring parameters from STPM0 to STPM6. Thus, if at least one of the monitoring parameters has a current value less than its associated limit, the regulation system, as well as, for example, the controller 75, detects operation of the power unit 2 during a risky flight phase.

[0093] One monitored parameter can be a function of the external temperature T0 of the air surrounding aircraft 1, and is measured using external temperature sensor 40 during step STPD0. The associated limit is then an external temperature limit LT0 stored in the regulation system 55 or even, for example, in the controller 75. The temperature value can then be equal to the external temperature T0, or calculated by the controller 75 using a stored law based on the external temperature T0 and the current speed of aircraft 1 measured using speed sensor 43. Detecting that the power unit 2 is in a risky flight phase of STPD then involves detecting, using the regulation system 55 or even the controller 75, that the STPD0 temperature value is less than the stored external temperature limit LT0.

[0094] One monitored parameter could be the oil temperature TOIL of the oil circulating in the passive engine, which is measured using oil temperature sensor 41 during step STPM1. The associated limit is then the oil temperature limit LTOIL stored in controller 75. Detecting STPD during a risky flight phase of the power unit 2 operation therefore involves using regulation system 55 or even controller 75 to detect that the STPC1 oil temperature TOIL has a current value lower than the stored oil temperature limit LTOIL.

[0095] One monitored parameter could be the fuel temperature TFUEL of the fuel to be supplied to the passive engine, which is measured using the fuel temperature sensor 42 in step STPM2. The associated limit is then stored in the controller 75 as a fuel temperature limit LTFUEL. Detecting STPM2 operation during a risky flight phase then involves using the regulation system 55 or even the controller 75 to detect that the STPM2 fuel temperature TFUEL is below the current value of the stored fuel temperature limit LTFUEL.

[0096] One monitored parameter could be the ambient temperature TCOMP in the engine compartment 100 housing the passive engine, measured using ambient temperature sensor 44 during step STPM3. The associated limit is the ambient temperature limit LTCOMP stored in controller 75. Detecting STPD operation of the power unit 2 during a risky flight phase then involves using regulation system 55 or even controller 75 to detect that the STPC3 ambient temperature TCOMP is below the current value of the stored ambient temperature limit LTCOMP.

[0097] One monitoring parameter may be the internal temperature LTENG in the passive engines 10, 11, measured using internal temperature sensor 45 during step STPM4. The associated limit is the internal temperature limit LTENG stored in controller 75. Detecting the operation of the power unit 2 during the risky flight phase of STPD then involves using regulation system 55 or even controller 75 to detect that the STPC4 internal temperature LTENG has a current value lower than the stored internal temperature limit LTENG.

[0098] One monitoring parameter could be the oil pressure POIL of the oil circulating in the passive engine, measured using oil pressure sensors 46, 460 in step STPM5. The associated limit is the oil pressure limit LPOIL stored in controller 75. Detecting STPD during a risky flight phase of power unit 2 operation then involves using regulation system 55 or even controller 75 to detect that the oil pressure POIL in STPC5 has a current value lower than the stored oil pressure limit LPOIL.

[0099] One monitored parameter could be the fuel pressure PFUEL, which supplies fuel to the internal combustion engine 10, measured using fuel pressure sensors 47, 470 in step STPM6. The associated limit is the fuel pressure limit LPFUEL stored in controller 75. Detecting STPD operation of the power unit 2 during a risky flight phase then involves using regulation system 55 or even controller 75 to detect that the STPC6 fuel pressure PFUEL has a current value lower than the stored fuel pressure limit LPFUEL.

[0100] Regardless of whether the aircraft 1 is in a risky flight phase, after detecting the operation of the power unit 2 in a risky flight phase, the controller 75 controls the exit from the asymmetric operation mode via the reactivation STPREAC of one or more passive engines.

[0101] According to the first variant, exiting the asymmetric operating mode involves sending an STPR1 alarm using siren 80. Controller 75 sends an alarm signal to siren 80, which then sounds an alarm. The pilot may, for example, use human-machine interface 88. During step STPR2, human-machine interface 88 sends an exit signal received by controller 75, or no longer sends a signal indicating the need to implement asymmetric mode. Controller 75 then controls the reactivation of STPR1 in a normal manner. For example, controller 75 transmits signals to the engine computer, which may optionally control the passive engine's starter motor and fuel metering device in a normal manner to enable cooperative operating mode.

[0102] According to the second variant, after detecting that the STPD is operating during a risky flight phase, the reactivation STPREC of the passive engine 11 is automatically executed by the regulation system 55. For example, the controller 75 transmits an output signal to the engine computer 61 of the passive engine 11, which can guide the starter motor and fuel metering valve of the passive engine in the normal manner.

[0103] Optionally, the pilot can use the human-machine interface 81 to select the variant to be applied during flight.

[0104] On the other hand, before initiating an exit from the asymmetric operating mode, controller 75 can be configured to detect whether at least one monitored parameter is greater than or equal to an associated limit and less than or equal to an associated threshold, wherein the associated threshold is greater than the associated limit. If this is the case, controller 75 can send an alarm signal to alarm 80 to generate an alarm signal indicating the entry into a transient phase approaching one or more limits requiring exit from the asymmetric operating mode.

[0105] Naturally, the invention can undergo many variations in its implementation. Although several embodiments have been described above, it should be readily understood that it is not conceivable to exhaustively characterize all possible embodiments. Of course, any of the described devices can be replaced with equivalent devices without departing from the scope of the invention and the claims.

Claims

1. A method for piloting an aircraft (1), the aircraft (1) having a power unit (2) comprising at least two internal combustion engines (10) and a transmission system (25) connected to at least one rotor (5), each internal combustion engine (10) having a power shaft (20) connected to the transmission system (25), the method comprising an asymmetric operating mode comprising using an adjustment system (55) to actively adjust at least one active engine of at least two internal combustion engines (10) at an active speed, the active engine outputting non-zero active drive power at an active speed and via its power shaft (20) to cause rotation of the rotor (5), the asymmetric operating mode comprising, together with adjusting the active engine at an active speed, placing at least one passive engine of at least two internal combustion engines (10) in a standby mode, the passive engine not transmitting power to the rotor (5), wherein during the asymmetric operating mode, the method comprises the following steps: ● Use the adjustment system (55) to detect (STPD) the operation of the power unit (2) in a risky flight phase that is not conducive to exiting the asymmetric operation mode before the flight phase; and ● use the adjustment system (55) to reactivate (STPREAC) the passive engine after detecting the operation of the power unit (2) in a risky flight phase.

2. The method according to claim 1, wherein, After detecting operation of the power unit (2) in a risky flight phase, the method includes issuing an alarm (STPR1) using an alarm device (80) and, after manipulation via the human-machine interface (88), reactivating (STPREAC) the passive engine by the regulation system (55).

3. The method according to claim 1, wherein, After the operation of the power unit (2) is detected to be in a risky flight phase (STPD), the passive engine is automatically reactivated (STPREAC) by the regulation system (55).

4. The method of claim 1, wherein the method comprises using various sensors (40 to 47, 410, 440, 450, 460, 470) of the regulation system (55) to measure (STPM01 to STPM6) at least one monitoring parameter, and the operation of detecting (STPD) the power unit (2) in a risky flight phase comprises detecting (STPC0 to STPC6) that the monitoring parameter has a current value less than an associated limit.

5. The method according to claim 4, wherein the at least one monitored parameter comprises at least one of the following parameters: • a temperature value, which is a function of the external temperature (T0) of the air surrounding the aircraft (1) as measured using an external temperature sensor (40), and the associated limit is a stored external temperature limit (LT0), and the operation of the power unit (2) in a risky flight phase includes detecting (STPC0) that the temperature value is less than the stored external temperature limit; • the oil temperature (TOIL) of the lubrication circuit of the passive engine as measured using an oil temperature sensor (41), and the associated limit is a stored oil temperature limit. The operation of the power unit (2) in the risky flight phase of detection (STPD) includes detecting (STPC1) that the oil temperature (TOIL) has a current value that is less than the stored oil temperature limit (LTOIL); and ● measuring the fuel temperature (TFUEL) of the fuel supplied to the passive engine using a fuel temperature sensor (42), and the associated limit is the stored fuel temperature limit (LTFUEL). The operation of the power unit (2) in the risky flight phase of detection (STPD) includes detecting (STPC2) that the fuel temperature (TFUEL) has a current value that is less than the stored fuel temperature limit (LTFUEL).

6. The method according to claim 5, wherein, The temperature value is equal to the external temperature (T0), or the temperature value is calculated based on the external temperature (T0) and the current speed of the aircraft (1) measured using the speed sensor (43).

7. The method of claim 4, wherein the at least one monitored parameter includes the ambient temperature (TCOMP) in the engine compartment (100) housing the passive engine, measured using an ambient temperature sensor (44), and the associated limit is a stored ambient temperature limit (LTCOMP), and the operation of detecting (STPD) the power unit (2) in the risky flight phase includes detecting (STPC3) that the ambient temperature (TCOMP) has a current value less than the stored ambient temperature limit (LTCOMP).

8. The method of claim 4, wherein the at least one monitoring parameter includes the internal temperature (LTENG) in the passive engine measured using an internal temperature sensor (45), and the associated limit is a stored internal temperature limit (LTENG), and the operation of detecting (STPD) the power unit (2) in a risky flight phase includes detecting (STPC4) that the internal temperature (LTENG) has a current value less than the stored internal temperature limit (LTENG).

9. The method according to claim 4, wherein, The at least one monitored parameter includes the oil pressure (POIL) of the lubrication circuit of the passive engine, which is measured using oil pressure sensors (46, 460), and the associated limit is a stored oil pressure limit (LPOIL). The operation of the power unit (2) in a risky flight phase includes detecting (STPC5) that the oil pressure (POIL) has a current value that is less than the stored oil pressure limit (LPOIL).

10. The method of claim 4, wherein the at least one monitored parameter includes fuel pressure (PFUEL) supplying fuel to the passive engine, and the associated limit is a stored fuel pressure limit (LPFUEL) measured using fuel pressure sensors (47, 470), and the operation of detecting (STPD) the power unit (2) in a risky flight phase includes detecting (STPC6) that the fuel pressure (PFUEL) has a current value less than the stored fuel pressure limit (LPFUEL).

11. The method of claim 4, wherein the method includes issuing an alarm using the alarm (80) when the at least one monitored parameter is greater than or equal to the associated limit and less than or equal to the associated threshold, the alarm signaling that the aircraft is in a transient phase of operation, the associated threshold being greater than the associated limit.

12. The method according to claim 1, wherein the operation of the power unit (2) in a risky flight phase that is not conducive to exiting the asymmetric operating mode is detected by the regulation system (55) before the flight phase, and is operated independently of the operation of at least one active engine of the at least two internal combustion engines (10).

13. An aircraft (1) having a power unit (2) comprising at least two internal combustion engines (10) and a transmission system (25) connected to at least one rotor (5), each internal combustion engine (10) having a power shaft (20) connected to the transmission system (25), the aircraft having an asymmetric operating mode comprising using an adjustment system (55) to actively adjust at least one active engine of the at least two internal combustion engines (10), the active engine outputting non-zero active drive power at an active speed and via its power shaft (20) to cause the rotor to rotate, the asymmetric operating mode comprising placing at least one passive engine of the at least two internal combustion engines in a standby mode together with adjusting the active engine at an active speed, the passive engine not transmitting any power to the rotor (5), wherein the adjustment system (55) is configured to implement the method according to claim 1.

14. The aircraft according to claim 13, wherein, The regulating system (55) includes at least one of the following sensors: an external temperature sensor (40) for measuring the external temperature (T0) of the air surrounding the aircraft (1); an oil temperature sensor (41, 410) for measuring the oil temperature (TOIL) of each internal combustion engine (10); a fuel temperature sensor (42) for measuring the fuel temperature (TFUEL) of the fuel supplied to each internal combustion engine (10); a speed sensor (43) for measuring the current speed of the aircraft (1); and a speed sensor for measuring the speed of the engine compartment (100) housing the internal combustion engines (10). Each internal combustion engine has an ambient temperature sensor (44, 440) for measuring the ambient temperature (TCOMP) in the internal combustion engine (10), an internal temperature sensor (45, 450) for measuring the internal temperature (LTENG) in the internal combustion engine (10), an oil pressure sensor (46, 460) for measuring the oil pressure in the internal combustion engine (10), and a fuel pressure sensor (47, 470) for measuring the fuel pressure of the fuel flowing in the internal combustion engine (10).

15. The aircraft according to claim 13, wherein the control system (55) includes an alarm (80).

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