Hybrid propulsion system for a helicopter

Abstract

A propulsion system (2) for a helicopter (1) is disclosed, comprising a main engine (9), a main rotor (3), a main gearbox (4) comprising an output mechanically connected to the main rotor (3), a reduction gearbox (13) mechanically coupled between the main engine (9) and a first input of the main gearbox (4), and an auxiliary device (10). The auxiliary device (10) comprises a first electric machine and a second electric machine, the first electric machine being mechanically coupled to the reduction gearbox (13) and being configured to act as a generator to take away energy produced by the main engine (9), the second electric machine being mechanically coupled to a second input of the main gearbox (4), the second electric machine being supplied with electric power by the first electric machine and being configured to act as an electric motor to deliver additional mechanical power to the main gearbox (4).

Description

Technical Field

[0001] This invention relates to the general field of helicopter propulsion systems, and more specifically to the field of hybrid propulsion systems for helicopters, particularly to helicopters referred to as single-engine helicopters. Background Technology

[0002] A single-engine helicopter is a helicopter that includes a propulsion system consisting of a single main engine, typically an internal combustion engine and, for example, a turbine engine, for driving the main rotor via a main gearbox called the MGB and the tail rotor (also known as the acronym ATR for anti-torque rotor) via a rear gearbox called the RGB.

[0003] The propulsion system may also include equipment for assisting the helicopter. This auxiliary equipment is used in emergencies to provide instantaneous power to the helicopter, and more precisely, to power the main rotor and rear rotor.

[0004] The first emergency is a main engine failure. In this situation, the pilot then initiates a degraded flight mode, known as autorotation. Auxiliary equipment allows for mechanical assistance to the helicopter during autorotation, particularly in the first and / or final phases of flight (to "pull-up" before landing). Therefore, this type of auxiliary equipment can significantly help the pilot complete an autorotation landing.

[0005] The second type of emergency is when additional power is needed immediately, for example, during obstacle avoidance or high-altitude temperature reversal.

[0006] Different architectures for integrating auxiliary equipment are known from document FR3 019 588. The auxiliary equipment includes a turbine powered by a gas generator with solid-state storage for driving shaft rotation, and controlled components for supplying power to the driving turbine. In this example, the mechanical rotational power of the shaft is used to drive the helicopter's main rotor by introducing this power directly at the MGB, or at the front drive shaft, or at the shaft of the free turbine of the turbine engine (main engine).

[0007] This type of propulsion system leads to size issues. In fact, integrating this type of auxiliary equipment into an already compact engine nacelle would require considerable changes to the helicopter's fuselage and to the power transmission between the main engine and the MGB.

[0008] Furthermore, introducing this power into the free turbine of the main turbine engine has several drawbacks.

[0009] The first drawback is that in the event of a gas supply failure to the free turbine of the main turbine engine, it will not generate driving torque and will decelerate very rapidly due to aerodynamic friction losses. These losses can reach tens of kilowatts (kW). Therefore, it is understandable that, depending on the application scenario mentioned above—whether the main engine fails or additional power is immediately needed to avoid an obstacle—the effective driving force seen by the main rotor will be different, which may surprise helicopter pilots.

[0010] The second drawback is that when auxiliary power is injected via the free turbine, it is impossible to supply such auxiliary power to the main gearbox and the rear gearbox in the event of a failure in the free turbine and / or the downstream mechanical components of the turbine engine (and especially the reduction gearbox of a turbine engine equipped with this type of reduction gearbox).

[0011] The third drawback is that a specific interface needs to be provided on the turbine engine to allow this auxiliary power to be injected into the free turbine.

[0012] Furthermore, the power of a helicopter's turbine engine is thermodynamically limited, and the power of the helicopter's turbine engine is also mechanically limited because the mechanical torque transmitted to the helicopter's main rotor passes through the reduction gearbox and the main gearbox.

[0013] At low ambient temperatures and altitudes, it is well known that the power delivered via the MGB input is designed to be limited by a torque value that cannot be continuously exceeded. Therefore, when the helicopter rotor speed decreases, for example to limit acoustic pollution, the power delivered to the rotor decreases proportionally. On the other hand, at higher temperatures and altitudes, the power delivered by the helicopter turbine engine is limited by the thermodynamic forces of the free turbine.

[0014] The propulsion system of a single-engine helicopter is known from document FR3 062 882, which includes a main engine connected to a front drive shaft and a rear drive shaft, which are capable of driving the MGB and RGB respectively; and an auxiliary device fixed to the main engine, which allows the RGB and MGB to be mechanically driven by introducing power on the rear drive shaft.

[0015] However, the auxiliary equipment described in this document uses auxiliary energy sources, such as pyrotechnics and / or electrical and / or hydraulic and / or pneumatic components. This auxiliary energy source forms an additional element to be integrated into the helicopter's propulsion system, thus increasing the volume and overall mass of the auxiliary equipment, and consequently increasing the volume and overall mass of the propulsion system. An auxiliary system with only a single reversible motor is also known from document FR3080 835, which needs to be associated with a battery to operate in motor mode and assist the main engine via the MGB. The availability and duration of the system's permissible assistance are necessarily limited by the battery capacity and its effective charging when needed. Summary of the Invention

[0016] To address this, the present invention proposes a helicopter propulsion system that includes auxiliary equipment that allows for the correction of the aforementioned drawbacks and, in particular, optimizes the integration of the auxiliary equipment while reducing mass and volume.

[0017] In one object of the present invention, a propulsion system for a helicopter is provided, comprising: a main engine, a main rotor, a main gearbox including an output mechanically connected to the main rotor, a reduction gearbox mechanically coupled between a first input of the main engine and the main gearbox, and auxiliary equipment.

[0018] According to the technical features of the present invention, the auxiliary equipment includes a first motor and a second motor, the first motor being mechanically coupled to a reduction gearbox and configured to operate as a generator to take energy generated by the main engine, the second motor being mechanically coupled to a second input of the main gearbox, the second motor being powered by the first motor and configured to operate as an electric motor to deliver additional mechanical power to the main gearbox.

[0019] Therefore, the propulsion system according to the invention proposes an architecture applicable to both turbine engines with free turbines and turbine engines with linked turbines, and allows the use of all thermodynamic power available upstream of the reduction gearbox via power take-off, which is achieved by means of two motors generating auxiliary coupling between the main engine and the main gearbox.

[0020] In fact, the propulsion system according to the invention thus allows for auxiliary equipment that uses the additional energy available from the main engine to compensate for the torque loss generated by the reduction gearbox. In other words, the power loss caused by the reduction gearbox is recovered due to the amount of energy available on the gas generator, and this energy can be converted into propulsion electrical energy by the auxiliary equipment, i.e., used to deliver additional mechanical power to the main gearbox.

[0021] Therefore, when the power delivered to the MGB is limited by the maximum torque allowed at the output of the engine propulsion unit, including the main engine and reduction gearbox, and / or at the input of the MGB, a motor mounted upstream of the reduction gearbox supplies electrical energy to another motor operating in motor mode at another input of the MGB. This additional power transmitted through the main rotor mast can be used in emergency situations requiring excess lift (to avoid obstacles, etc.) without damaging the reduction gear.

[0022] According to a first aspect of the propulsion system, the propulsion system may further include an electronic control unit electrically coupled to the first motor and the second motor and configured to control the operation of the first motor and the second motor according to the available output torque of the main engine and the output torque required by the main gearbox.

[0023] According to a second aspect of the propulsion system, the electronic control unit may include a measuring component and a control component. The measuring component is configured to continuously measure the instantaneous torque at a first input of the main gearbox, and the control component is configured to control the operation of the first motor to supply power to the second motor when the torque measured by the measuring component is less than the output torque required by the main gearbox, and to command an increase in operating speed to supply additional energy required for the electrical power supply of the second motor.

[0024] The instantaneous torque measured at the first input of the main gearbox can also be compared with the maximum permissible torque at that first input. If the measured torque reaches the maximum permissible torque, the electronic control unit can command that the operating speed of the main engine should not be increased, while continuing to command the first motor to operate as a generator to supply power to the second motor.

[0025] According to a third aspect of the propulsion system, the control unit can be configured to control the first and second motors to compensate for insufficient output power of the main gearbox when the main rotor operates at a speed lower than the nominal speed during a flight phase in which the main rotor needs to operate at its nominal speed.

[0026] According to a fourth aspect of the propulsion system, the main engine may include an output shaft mechanically connected to a reduction gearbox, and a first motor includes a rotor and a stator, the rotor being formed by an electromagnetic portion of the main engine's output shaft, and the stator being mounted around said electromagnetic portion of the main engine's output shaft.

[0027] According to the fifth aspect of the propulsion system, in the first motor, the reduction gearbox can be a planetary gear reduction gearbox that is simultaneously coupled to the main engine, the first motor and the main gearbox.

[0028] According to a sixth aspect of the propulsion system, the propulsion system may further include an overrunning clutch coupled to the output of the reduction gear, a rear drive shaft coupled between the overrunning clutch and the anti-torque rotor, and a front drive shaft coupled between the first input of the main gearbox and the overrunning clutch.

[0029] According to the seventh aspect of the propulsion system, the auxiliary equipment may further include at least one component electrically coupled to the first motor for storing energy. The electronic control unit can then be configured to transfer electrical energy from the first motor to the energy storage device or to the second motor, based on the energy demand while the first motor is generating electricity.

[0030] According to another objective of the present invention, a helicopter comprising a propulsion system as defined above is proposed.

[0031] According to another object of the invention, a method is proposed for assisting a propulsion system of a helicopter as defined above, the method comprising: requesting activation to draw mechanical energy from the main engine of the propulsion system by means of a command to operate a first motor as a generator, the activation request being the result of a decision made by at least one specific condition verified by the helicopter pilot, or by the engine control computer, or by the helicopter's avionics equipment.

[0032] In one aspect of the auxiliary method, the activation requirement may be the result of a decision by the engine's control computer when it detects that the main engine's operating point is located within a set of predetermined operating points for which the main engine's operating speed is below a percentage threshold of the main engine's nominal operating speed, and an acceleration transient returning to a higher speed above the percentage threshold is considered too long compared to the desired dynamics of the main engine and the helicopter, wherein the activation requirement is accompanied by a command to increase the main engine's operating speed to reduce the duration of possible acceleration transients to correspond to the desired dynamics.

[0033] Preferably, the percentage threshold of the nominal operating speed of the main engine is between 80% and 85%. Attached Figure Description

[0034] [ Figure 1 ] Figure 1 This is a schematic diagram of a helicopter equipped with a propulsion system according to a first embodiment of the present invention.

[0035] [ Figure 2 ] Figure 2 This is a schematic diagram of a helicopter equipped with a propulsion system according to a second embodiment of the present invention.

[0036] [ Figure 3 ] Figure 3A flowchart is shown of a method for an auxiliary helicopter propulsion system according to one embodiment of the present invention. Detailed Implementation

[0037] exist Figure 1 The diagram schematically illustrates a helicopter 1 including a propulsion system 2 according to a first embodiment of the invention.

[0038] The propulsion system 2 includes a main rotor 3, a main gearbox 4 (hereinafter referred to as MGB), a front drive shaft 5, a rear gearbox 6 (hereinafter referred to as RGB), and a rear rotor 7, commonly known as an anti-torque rotor (ATR). The main rotor 3 is driven by MGB 4, which itself is driven by the front drive shaft 5. Similarly, the rear rotor 7 is driven by RGB 6, which itself is driven by the rear drive shaft 8. In this example, the front and rear drive shafts 5 and 8 are substantially coaxial.

[0039] The propulsion system 2 of helicopter 1 also includes the main engine 9 and auxiliary equipment 10 for use in emergency situations, so as to provide additional power to helicopter 1 instantly, and more precisely, to the main rotor 3.

[0040] according to Figure 1 and Figure 3 In the embodiment illustrated, the main engine 9 includes an output shaft 12, which is mechanically connected to the front and rear drive shafts 5 and 8 via a reduction gearbox 13 and then a first overrunning clutch 14 (referred to as a propulsion overrunning clutch). It should be noted that in this invention, the terms "front" and "rear" associated with the drive shafts 5 and 8 are expressed in relation to the first overrunning clutch 14.

[0041] like Figure 1 As illustrated, the main engine here is a turbine engine formed by a gas generator 15 and a free turbine 16, with an output shaft 12 attached to the free turbine 16. The gas generator 15 includes, in a known manner, at least one air compressor 17, which supplies fuel in compressed air to a combustion chamber 18. The compressed air delivers hot gas to at least one gas expansion turbine 19, which drives the compressor 17 to rotate via a drive shaft 20. The gas then drives the free turbine 16, also known as the power turbine, for transmitting power.

[0042] As a variant, the main engine 9 can typically correspond to any type of internal combustion engine.

[0043] like Figure 1 As shown in the figure, the auxiliary equipment 10 includes a first motor 21, a second motor 22, and an electronic control unit 23.

[0044] The first motor 21 is mechanically coupled to the reduction gearbox 13 and is configured to operate as an electric power generator, thereby converting the mechanical energy taken off from the main engine 9 via the reduction gearbox 13 into electrical energy.

[0045] The second motor 22 may be mechanically connected to MGB 4 via gear train 24, such as Figure 1 As illustrated in the diagram, it is electrically connected to the first motor 21 via at least one cable 25. Due to the cable 25, the first motor 21 can use a portion of the energy delivered by the main motor 9 and recovered via the reduction gear 13 to supply power to the second motor 22.

[0046] The second motor 22 is configured to operate as an electric motor in itself, and thus uses electrical energy supplied by the first motor via cable 15 to deliver mechanical energy to MGB 4 in the form of mechanical torque.

[0047] The electronic control unit 23 is coupled not only to the first motor 21 and the second motor 22, such as Figure 1 As shown in the diagram, it is also coupled to the main engine 9.

[0048] The electronic control unit 23 is configured to control the operation of the first motor 21 and the second motor 22 based on the torque received by the main rotor 3 at the input of the MGB 4 and the torque required at the output of the MGB 4.

[0049] MGB 4 includes a first input 41 mechanically connected to the front drive shaft 5 and a second input possibly mechanically connected to the second motor 22 via a gear train 24. The electronic control unit 23 includes a torque measurement module 26 mounted on the first input 41 and configured to continuously measure the torque received at the first input 41 of the MGB, which is mechanically derived from the reduction gearbox 13 via an overrunning clutch 14.

[0050] When the output torque required by MGB 4 is greater than the torque measured at its first input 41, the electronic control unit 23 controls the main engine 9 to generate a greater mechanical torque than it currently produces, and activates the first motor 21 so that it takes away the additional torque supplied by the main engine 9, and thus supplies power to the second motor 22, which can then supply additional mechanical torque to the second input 42 of MGB 4. MGB 4 can then supply torque greater than, or even equal to, the mechanical torque received at its first input 41, the torque required by the main rotor 2.

[0051] Therefore, when the speed of the gas generator 15 (NG) of the turbine engine is low (i.e., less than 85% of the nominal speed of the gas generator 15), the transient return to high speed (i.e. greater than 85% of the nominal speed of the gas generator) using the auxiliary equipment described in the prior art may be long, and the dynamics of the turbine engine and thus the dynamics of the helicopter are severely affected.

[0052] Since the power take-off is achieved by the auxiliary device 10 of the propulsion system 2 according to the invention, and since the main engine 9, the first motor 21, and the second motor 22 are controlled by the electronic control unit 23, it is possible to manually take off the power through the circuit at the said high speed.

[0053] By maintaining a high speed, a portion of the power can be allocated to the mechanical torque injected into MGB 4, and the remainder to the first motor 21 that will generate electricity, which can optionally power the second motor or charge an energy storage device (such as a battery). It is also conceivable to provide components for dissipating the electricity generated by the first motor 21, such as by means of heat dissipation via resistors, which can be selectively controlled, particularly in cases where it is desired to artificially increase the speed of the gas generator, as mentioned below. This dissipation of the generated electricity can allow for an increase in the energy taken off by power takeoff, and / or dissipation of energy taken off when the battery is full and the second motor 22 does not need to be activated, and thus, in any case, increase the speed of the gas generator while still maintaining the same mechanical torque injected into MGB 4.

[0054] To activate the power-driven removal (step 340), at least two conditions must be combined, such as... Figure 3 As indicated in the document, Figure 3 A flowchart of a method for a propulsion system for an auxiliary helicopter according to one embodiment is presented. Following the activation requirement of power take-off (step 300), it is necessary to verify (step 320) that the limit of the drive torque is occurring, which corresponds to the limit of the torque that the reduction gear can transmit and / or the MGB can accept at its first input 41, and to verify (step 330) that the thermal limit of the engine has not occurred (this limit may depend on the speed of the gas generator 15 or on the temperature in a given compartment of the engine).

[0055] Power take-off activation requests may be the result of decisions made based on at least one specific condition, which is verified by the helicopter pilot, the engine control computer (such as FADEC), or even the helicopter's avionics; this list is not exhaustive. Therefore, among the various types of decisions that can trigger power take-off activation requests, the following three categories will be of particular attention.

[0056] This decision stems from the pilot's understanding. The pilot knows that by requesting power take-off activation at the limits of the gas turbine's output torque (or at the MGB input), he will have a more responsive propulsion system. In fact, power take-off activation results in an increase in the operating speed of the propulsion system's main engines—that is, the speed of the engine's gas generator, or, in the case of a twin-engine or tri-engine helicopter, the speed of the main engine. This increased speed benefits the propulsion system's acceleration capabilities, particularly during takeoffs without engine failure (for all engines operating, known as AEO takeoff), but especially during takeoffs when the helicopter's propulsion system has two or three engines and one engine fails (for one engine not operating, known as OEI takeoff).

[0057] The decision originates from the engine control computer (FADEC). In the first configuration, the computer monitors the main rotor speed error in real time; this error is the difference between the measured speed of the main rotor and its setpoint speed. If this difference is too high (the difference between the setpoint and the measured value exceeds a threshold), it indicates a large transient additional power demand. Power take-off can then help increase the effective speed of the main rotor.

[0058] In the second configuration, the computer monitors the main rotor speed in real time over time. There is no need to wait for the main rotor speed to reach a critical value, such as below 85% of its nominal speed; if the descent slope of the main rotor over time exceeds an absolute threshold, the computer triggers power take-off.

[0059] The decision stems from avionics. For example, if the helicopter's horizontal speed, as measured by avionics, is less than a threshold, such as 50 knots, it indicates that the helicopter is in the takeoff or landing phase. Therefore, it is advantageous to increase the gas generator speed by removing power at the limit of the free turbine's output torque to compensate for possible takeoff or landing in the event of one engine failure (OEI). In another example, if the main rotor speed setpoint is stabilized below the main rotor's nominal speed, such as 90% of the nominal speed, to limit noise, this means that the main rotor speed is regulated to a lower value relatively close to the main rotor's stall speed (which is, for example, between 80% and 85% of the nominal speed). Power removal then allows for an increase in the NG speed of both main engines in a twin-engine propulsion system. This means that in the event of an OEI failure in one engine, the higher speed of the healthy engine will allow it to have an additional torque reserve, which can be immediately supplied to the main rotor by disabling power removal, thus allowing for a limitation on the NR speed drop after an OEI failure, keeping the NR speed above the stall speed.

[0060] Once power take-off is activated, the efficiency degradation at certain operating points will certainly increase fuel consumption slightly, but in particular, it will allow for an artificial increase in the speed of gas generator 15 while maintaining the same mechanical torque injected into MGB 4.

[0061] By storing the operating points of the main engine 9 in the memory of the electronic control unit where unsatisfactory transients during acceleration are not permitted, thus ensuring the proper operation of the main rotor 3 of the helicopter 1, the auxiliary equipment 10 allows for an available instantaneous reserve of additional torque. Therefore, when the engine control computer detects that the main engine's operating point is within a set of predetermined operating points, the activation request for power take-off may be a result of a decision by the engine control computer for these predetermined operating points where the main engine's operating speed is below a percentage threshold of the main engine's nominal operating speed, and an acceleration transient returning to a higher speed above said percentage threshold is considered too long compared to the desired dynamics of the main engine and therefore the desired dynamics of the helicopter. The activation request for power take-off is then accompanied by a command to increase the NG operating speed of the main engine, thereby reducing the duration of possible acceleration transients to correspond to the desired dynamics of the main engine.

[0062] The percentage threshold for the operating speed of the main engine can be included between 80% and 85%.

[0063] This control strategy and the expected results are possible because the speed of electrical technology and control causes an instantaneous drop in torque on the output shaft of the turbine engine in response to power loss.

[0064] Figure 2 The diagram schematically illustrates a helicopter 1 equipped with a propulsion system 2 according to the second embodiment.

[0065] and Figure 1 The same elements in the first embodiment illustrated herein have the same numerical designations.

[0066] Figure 2 The second embodiment illustrated in the figure is the same as Figure 1 The difference in the first embodiment illustrated is that the first motor 21' of the auxiliary device 10' includes a rotor formed by a portion of the output shaft 12 of the main engine 9 (and more particularly by electromagnetic elements mounted on the mechanical output shaft 12), and a stator 210 arranged around the portion of the output shaft 12 that forms the rotor.

[0067] The first motor 21' is therefore connected upstream of the reduction gear 13, more specifically between the main engine 9 and the reduction gearbox 13.

[0068] Therefore, the propulsion system according to the present invention provides an auxiliary device that allows for optimized integration of the auxiliary device by reducing mass and volume.

Claims

1. A propulsion system (2) for a helicopter (1) includes a main engine (9), a main rotor (3), an electronic control unit (23), a main gearbox (4) including an output mechanically connected to the main rotor (3), a reduction gearbox (13) mechanically coupled between the main engine (9) and a first input (41) of the main gearbox (4), and auxiliary equipment (10). in, The auxiliary equipment (10) includes a first motor (21') and a second motor (22). The first motor (21') is mechanically coupled to a reduction gearbox (13) and configured to operate as a generator to draw energy from the main engine (9). The second motor (22) is mechanically coupled to a second input (42) of the main gearbox (4) via a gear train. The second motor (22) is powered by the first motor (21') and configured to operate as an electric motor to deliver additional mechanical power to the main gearbox (4) when needed. Specifically, the mechanical coupling from the first motor to the reduction gearbox is independent of the mechanical coupling from the reduction gearbox to the main engine. The main engine (9) includes an output shaft (12) driven by a power turbine (16) of the main engine and mechanically connected to a reduction gearbox (13), and a first motor (21') includes a rotor and a stator (210), the rotor being formed by an electromagnetic portion of the output shaft (12) of the main engine (9), and the stator (210) being mounted around the electromagnetic portion of the output shaft of the main engine (9), and, The electronic control unit (23) is configured to control the output torque of the main engine.

2. The propulsion system (2) according to claim 1, wherein the electronic control unit (23) is electrically coupled to the first motor and the second motor (21', 22) and configured to control the operation of the first motor and the second motor (21', 22) according to the available output torque of the main engine (9) and the output torque required by the main gearbox (4).

3. The propulsion system (2) according to claim 2, wherein the electronic control unit (23) includes a measuring component (26) configured to continuously measure the instantaneous torque at a first input (41) of the main gearbox (4), and the electronic control unit is configured to control the operation of the first motor (21, 21') to supply power to the second motor (22) when the instantaneous torque measured by the measuring component (26) is less than the required output torque of the main gearbox (4), and to command an increase in the operating speed of the main engine (9) to supply additional energy required for the power supply of the second motor (22).

4. The propulsion system (2) according to claim 2, wherein the electronic control unit (23) is configured to control the operation of the first motor and the second motor (21', 22) to compensate for insufficient output power of the main gearbox (4) when the main rotor (3) operates at a speed lower than the nominal speed during a flight phase in which the main rotor (3) needs to operate at its nominal speed.

5. The propulsion system (2) according to claim 1, wherein the reduction gearbox (13) is a planetary reduction gearbox that is simultaneously coupled to the main engine (9), the first motor (21') and the main gearbox (4).

6. The propulsion system (2) according to claim 1 further includes an overrunning clutch (14) coupled to the output of the reduction gearbox (13), a rear drive shaft (8) coupled between the overrunning clutch (14) and the anti-torque rotor (7), and a front drive shaft (5) coupled between the first input (41) of the main gearbox (4) and the overrunning clutch (14).

7. A helicopter comprising the propulsion system (2) according to claim 1.

8. A method for assisting a system for a propulsion system (2) of a helicopter (1) according to claim 1, the method comprising: The activation of the main engine (9) of the propulsion system (2) is requested by means of the command to operate the first motor (21') as a generator. The activation request is the result of a decision made by at least one specific condition verified by the helicopter pilot, or by the main engine control computer, or by the helicopter's avionics.

9. The auxiliary method according to claim 8, wherein the activation requirement is the result of a decision by the main engine control computer when it detects that the operating point of the main engine (9) is within a set of predetermined operating points, for which the operating state of the main engine is below a percentage threshold of the nominal operating speed of the main engine, and an acceleration transient returning to a faster speed greater than the percentage threshold is considered too long compared to the desired dynamics of the main engine and the helicopter, and wherein the activation requirement is accompanied by a command for increasing the operating speed of the main engine to reduce the duration of possible acceleration transients to correspond to the desired dynamics.

10. The auxiliary method of claim 9, wherein the percentage threshold of the nominal operating speed of the main engine is between 80% and 85%.

Images