Method for assisting propulsion by detecting a failure of an aircraft turbine engine

Abstract

This method of assisting propulsion by detecting a failure of a turbine engine of an aircraft operating in a nominal mode, the aircraft comprising a deactivated assistance engine, comprises a step (2) of comparing operating parameters of the turbine engine with the equivalent parameters of a model representative of a healthy turbine engine, a step (4) of detecting a failure of the turbine engine by detecting an anomaly of at least one operating parameter of the turbine engine, a step (6) of choosing an activation mode of the assistance engine as a function of the operating parameters of the turbine engine and / or flight parameters of the aircraft, and a step (8) of activating the assistance engine with the chosen activation mode. Figure for the abstract: Fig. 1

Description

Title of the invention: Method for assisting propulsion by detecting a failure of an aircraft turbine engine Technical field

[0001] The present invention relates to the assistance of an aircraft turbomachine following the detection of a failure in said aircraft.

[0002] In particular, the present invention relates to helicopters with twin-engine architecture in which one of the two engines can operate in a deactivated or energy-saving mode, also called standby mode.

[0003] Generally speaking, the invention applies to all aircraft comprising at least two engines. Previous techniques

[0004] A turbomachine, and in particular a turbomachine of an aircraft, for example a helicopter turboshaft, must be started before it can provide thrust power or mechanical power to the rotor. In particular, the start-up includes heating the components of the turbomachine as well as starting the turbine to rotate.

[0005] During a flight phase of an aircraft comprising a two-turboengine architecture, one turboengine may sometimes be switched off due to the implementation of a particular flight mode, for example an energy saving mode, also called standby mode, the other turboengine operating in nominal mode.

[0006] In this standby mode, also detailed in document FR2967132A1, the turbine engine remains rotating at a low speed while being driven by combustion gases or by an assistance device such as an electric machine. In this standby mode, the combustion chamber can be switched off. Alternatively, the combustion chamber is switched on while the turbine engine is or is not assisted in its rotation.

[0007] To exit this standby mode, particularly in the event of a failure of the turbine engine operating in nominal mode, it is possible to restart the turbine engine using a conventional restart or even with a rapid restart. Document FR3027058A1 mentions in particular that a normal restart takes place over a period of 10 seconds to 1 minute while a rapid restart takes place over a period of between 5 and 15 seconds.

[0008] In order to detect failures or breakdowns in a turboshaft engine of a twin-turboshaft architecture, it is customary to compare the performance and parameters of the two turboshaft engines when the latter operate in the same manner. at the same rate.

[0009] However, when one of the two turboshaft engines is in a standby mode, it is not possible to detect failures by comparing the two turboshaft engines and the consistency of their performance and their operating parameters, the rotation speed and the temperature of the turboshaft engine in standby being lower than those of the turboshaft engine in nominal mode. Statement of the invention

[0010] The present invention therefore aims to overcome the aforementioned drawbacks and to provide improved detection of failures of a turbine engine, as well as better management of the activation of an assistance motor when a failure is detected.

[0011] The present invention relates to a method for assisting propulsion by detecting a failure of a turbine engine of an aircraft operating in a nominal mode, the aircraft comprising a deactivated assist engine. Nominal mode is understood to mean the normal operating mode where the engine provides power. The method comprises a step of comparing operating parameters of the turbine engine with the equivalent parameters of a model representative of a healthy turbine engine, a step of detecting a failure of the turbine engine by detecting an anomaly in at least one operating parameter of the turbine engine, a step of choosing an activation mode of the assist engine as a function of the operating parameters of the turbine engine and / or flight parameters of the aircraft, and a step of activating the assist engine with the chosen activation mode,the operating parameters of the turbine engine or detectable failures comprising at least one of the following parameters: a deterioration in the quality of regulation of the performance of the turbine engine, an incompatibility between the nominal operating mode of the turbine engine and reduced margins with respect to the operating limits of the rotational speed of the turbine engine, the internal temperature of the turbine engine, the torque supplied by the turbine engine, the acceleration or deceleration of said turbine engine, or the fuel flow limits, an abnormal vibration level of the turbine engine, an abnormally high temperature of an aircraft computer, a connection failure between an electronic control system of the turbine engine and a central computer of the aircraft, a detection of a fire in the turbine engine, a detection of a power oscillation of the turbine engine.

[0012] Thus, the detection of failures on a turbine engine is not carried out in comparison with another active turbine engine but in comparison with an internal model representative of a healthy turbine engine. In the present invention, the operating parameters of the turbine engine and the detectable failures have a definition equivalent when the operating parameters of the turbine engine are faulty.

[0013] In one embodiment, the flight parameters of the aircraft include the flight altitude of the aircraft, the choice of the activation mode of the assistance engine being made as a function of the possibility of maintaining the flight altitude of the aircraft for the time necessary for the activation of the assistance engine.

[0014] Advantageously, the step of comparing operating parameters of the turbine engine with the equivalent parameters of a model representative of a healthy turbine engine is carried out by a computer of the turbine engine storing the model representative of a healthy turbine engine.

[0015] In one embodiment, the step of detecting a failure of the turbine engine comprises measuring a significant difference between an operating parameter of the turbine engine and the equivalent parameter of the model representative of a healthy turbine engine.

[0016] Advantageously, the choice of the activation mode of the assistance motor is made between a normal activation mode having a first activation duration, an accelerated activation mode having a second activation duration and a rapid activation mode having a third activation duration, the first duration being longer than the second duration, and the second duration being longer than the third duration, the normal activation mode being favored over the accelerated activation mode, the accelerated activation mode also being favored over the rapid activation mode.

[0017] In a particular embodiment, the assistance engine is a second turboshaft engine.

[0018] Advantageously, the normal activation mode comprises a thermal stabilization step at a predefined temperature of the assistance engine and is preferred when no failure of the turbine engine is detected and an on-board computer and / or a pilot of the aircraft gives the instruction to the assistance engine.

[0019] Advantageously, the operating parameters of the turbine engine or the detectable failures include the fuel flow rate, and / or the temperature and / or the rotational speed at the outlet of a high-pressure turbine of the turbine engine, and / or the pressure at the outlet of a compressor of the turbine engine, and / or the torque, and / or the extinction of a combustion chamber of the turbine engine, and / or a leak on an oil circuit, and / or a leak on a fuel circuit, and / or a detection of filings, and / or an inability to regulate the performance of the turbine engine, and / or an inconsistency between the power demand and the power supplied, and / or a potential impact on the trajectory of the aircraft.

[0020] In one embodiment, the assist motor is a second turboshaft engine, and the normal activation mode, the accelerated activation mode and the mode rapid activation mode includes assistance of the second turbine engine by an electrical machine of the aircraft, the normal activation mode includes thermal stabilization at a predefined temperature of the second turbine engine and the implementation of a fuel flow law allowing an activation time of between 1 and 3 minutes, the accelerated activation mode includes the implementation of a fuel flow law allowing an activation time of between 10 seconds and 1 minute, and the rapid activation mode includes the use of a specific starting system configured to implement an assistance torque and a fuel flow law allowing an activation time of the second turbine engine of between 5 seconds and 15 seconds.

[0021] In a particular mode of implementation, the deactivated assistance motor corresponds to an operating mode of the assistance motor equivalent to a standby mode or to an intermediate mode between a standby mode and a nominal operating mode. Brief description of the drawings

[0022] Other aims, characteristics and advantages of the invention will appear on reading the following description, given solely by way of non-limiting example, and made with reference to the appended drawings in which:

[0023] [Fig.l] illustrates the steps of the method according to the invention;

[0024] [Fig.2] is a block diagram of the different states of the assistance motor depending on the activation mode chosen for said assistance motor during the implementation of the method according to the invention.

[0025] Detailed description of at least one embodiment

[0026] In one embodiment, the method according to the invention is implemented in an aircraft, for example a helicopter, comprising a turbine engine and an assist engine capable of supplying power to a main transmission gearbox of the aircraft.

[0027] The assist motor is for example an electric machine or a second turboshaft engine. The present description will be made with reference to a second turboshaft engine, the turboshaft engine previously mentioned being considered as a first turboshaft engine.

[0028] The first and second turboshaft engines each comprise, for example, a compressor which raises the pressure of a gas entering the turboshaft engines, a combustion chamber which raises the temperature of the compressed gas, and an expansion turbine which also drives the compressor.

[0029] In particular, the first and second turboshaft engines are, for example, free turbine turboshaft engines, the expansion turbine being divided into a high pressure turbine, also called a gas generator turbine, and a power turbine also called a free turbine.

[0030] The method according to the invention is in particular implemented when the aircraft operates with a specific consumption mode such as an energy saving mode, in particular when one of the two turbine engines, for example the first turbine engine, operates in a nominal mode, in other words in an operating mode offering standard performances and allowing the aircraft to move in the atmosphere, and when the second turbine engine operates in a deactivated mode. In particular, deactivated mode means a standby mode of the second turbine engine, or a consumption mode intermediate to the standby mode and the nominal mode of the second turbine engine.

[0031] The intermediate consumption mode is for example an operating mode of the second turbine engine when the latter switches from its nominal mode to its standby mode, or from its standby mode to its nominal mode.

[0032] The standby mode is a consumption mode allowing the second turbine engine of the aircraft to have low fuel consumption, while allowing activation which can be rapid, for a situation where the aircraft would need both turbine engines or in the event where the first turbine engine had a failure.

[0033] [Fig.l] shows schematically the different stages of the method of assisting propulsion by detecting a failure of an aircraft turbine engine.

[0034] In implementing this method, a step 2 is first carried out of comparing the operating parameters of the first turbine engine with the equivalent parameters of a model representative of a healthy turbine engine.

[0035] The representative model of a healthy turbine engine is for example stored in a computer of the first turbine engine, said computer implementing step 2 of comparing the operating parameters.

[0036] A step 4 is then carried out for detecting an anomaly in at least one of the operating parameters of the first turbine engine.

[0037] This step 4 of detecting an anomaly comprises, for example, the measurement of a significant difference between an operating parameter of the first turbine engine and the equivalent parameter of the model representative of a healthy turbine engine. A significant difference is understood to mean a difference whose value is predetermined for each operating parameter and symptomatic of a failure in the first turbine engine.

[0038] The operating parameters of the turboshaft engine or detectable failures include at least one of the following parameters: a deterioration in the quality of regulation of the performance of the turbine engine, an incompatibility between the nominal operating mode of the turbine engine, or more broadly the energy saving mode of the aircraft, and reduced margins compared to the operating limits of the rotation speed of the turbine engine, the internal temperature of the turbine engine, the torque supplied by the turbine engine, the acceleration or deceleration of said turbine engine, or the fuel flow limits, an abnormal vibration level of the turbine engine, an abnormally high temperature of a computer of the aircraft, a connection failure between an electronic control system of the turbine engine, also known in the prior art as FADEC, and a central computer of the aircraft and / or of the second turbine engine, a detection of a fire in the turbine engine, a detection of a power oscillation of the turbine engine.

[0039] Furthermore, the operating parameters of the first turbine engine and, more broadly, the detectable anomalies include anomalies concerning the fuel flow rate, and / or the temperature at the outlet of the high-pressure turbine and / or the rotational speed at the outlet of a high-pressure turbine of the turbine engine, and / or the pressure at the outlet of a compressor of the turbine engine, and / or the torque, and / or the extinction of a combustion chamber of the turbine engine, and / or a leak in an oil circuit, and / or a leak in a fuel circuit, and / or a pressure difference between a predetermined pressure and the pressure present in the oil or fuel circuits, and / or a detection of filings, and / or the activation of overspeed protection, and / or an inability to regulate the performance of the turbine engine, and / or an inconsistency between the power demand and the power supplied, and / or a potential impact on the trajectory of the aircraft,for example detected via an anomaly in the rotation speed of the free turbine of the turboshaft engine, indicating over-revving or under-revving, via a variation in atmospheric pressure indicating the aircraft moving away from or approaching the ground, via a variation in the power demand, or via a variation in the power consumed by the engine, or via the reaching of a minimum or maximum power threshold.

[0040] A step 6 is then carried out for choosing an activation mode for the second turbine engine as a function of the previously detected failure, and more broadly as a function of the operating parameters of the first turbine engine and / or the flight parameters of the aircraft, and / or the potential impact on the trajectory of the aircraft.

[0041] Finally, once the activation mode has been chosen, a step 8 of activating the second turboshaft engine is carried out according to said chosen activation mode.

[0042] [Fig.2] shows schematically a block diagram of the different states of the second turboshaft engine according to the activation mode chosen for the second turboshaft engine.

[0043] [Fig. 2] shows in particular a nominal consumption mode 10 of the second turbine engine, a standby mode 12, as well as an intermediate consumption mode 14 in the standby mode 12 and in the nominal mode 10. The step 8 of activating the second turbine engine according to the activation mode chosen applies to the second turbine engine in particular when the latter is in an operating mode such as the standby mode 12 or the intermediate consumption mode 14. In these latter modes, the combustion chamber can be switched off, while the turbine of the gas generator is weakly driven in rotation, for example using an electric machine. Alternatively, the combustion chamber is switched on and the turbine of the gas generator includes or does not include rotational drive assistance.

[0044] During step 8, it is possible to implement three different activation modes, the choice among these modes being made automatically by the computer of the second turboshaft engine or manually by a pilot of the aircraft.

[0045] In particular, the three activation modes can have a different impact on the aging of the second turbine engine.

[0046] Among the three possible modes, it is possible to carry out a step of implementing a normal activation mode 16 having a first duration of the second turbine engine in the nominal mode of said second turbine engine when the aircraft and / or the first turbine engine and / or the pilot verifies a first condition CL. This is the longest possible activation mode, it makes it possible in particular to preserve the mechanical integrity of the second turbine engine.

[0047] In one embodiment, the first condition C1 comprises the reception by the second turbine engine of an activation instruction sent by an on-board computer and / or by a pilot of the aircraft, and / or by the FADEC of the second turbine engine by means of an internal condition.

[0048] This normal activation mode 16 comprises a step of thermal stabilization at a predefined temperature of the second turbine engine. Alternatively, this mode 16 comprises a step of progressive powering up of the second turbine engine until the power supplied between the first and second turbine engines is balanced.

[0049] The normal activation mode 16 further comprises the implementation of a fuel flow law and torque optimization allowing an activation time of between 1 and 3 minutes.

[0050] This mode 16 thus makes it possible to gradually change the temperature of the components of the second turboshaft engine.

[0051] This normal activation mode 16 is the most commonly used activation mode since it is often a non-urgent activation. This is the mode with the least impact on the aging of the turboshaft engine and is therefore preferred.

[0052] The second type of possible activation is the implementation of an accelerated activation mode 18 having a second duration of the second turbine engine in the nominal mode of said second turbine engine, and this when the aircraft and / or the first turbine engine verifies a second condition C2.

[0053] The second condition C2 comprises, for example, the detection of a failure on the first and / or the second turbine engine and / or on the rest of the aircraft, and more broadly of an anomaly reflecting a failure, for example the loss of redundancy of a sensor, an excessive oil temperature, or the loss of an energy source of the aircraft.

[0054] Alternatively, the second condition C2 comprises the detection of flight conditions requiring operation in nominal mode 10 of the second turbine engine. For example, the measured atmospheric pressure is not compatible with the use of a single turbine engine in nominal mode, and it is therefore necessary to activate the second turbine engine. Alternatively, the second condition C2 comprises the reception by the second turbine engine of an activation instruction issued by an on-board computer and / or a pilot of the aircraft.

[0055] The accelerated activation mode 18 is substantially similar to the normal activation mode 16, but includes the implementation of a fuel flow law allowing an activation time of between 10 seconds and one minute. This accelerated activation mode 18 is therefore shorter, in particular thanks to the absence of a thermal stabilization step and allows the second turbine engine to be activated quickly. This activation nevertheless has a strong impact on the aging of the second turbine engine.

[0056] The third activation mode is the implementation of a rapid activation mode 20 having a third duration of the second turbine engine in the nominal mode 10 of said second turbine engine when the aircraft and / or the first turbine engine verifies a third condition C3.

[0057] The third condition C3 also comprises the detection of a failure, for example on the first turbine engine, or more broadly of an anomaly reflecting a fault generating a loss of power, for example an anomaly among the values ​​of fuel flow rate, and / or of the temperature and / or of the rotational speed at the outlet of a high-pressure turbine of the first turbine engine, and / or of the pressure at the outlet of a compressor of the turbine engine, and / or of the torque, and / or the detection of the extinction of a combustion chamber of the turbine engine, and / or of a leak on an oil circuit, and / or of a leak on a fuel circuit, and / or a detection of filings, and / or of an inability to regulate the performance of the turbine engine, and / or of an inconsistency between the power demand and the power supplied, and / or a potential impact on the trajectory of the aircraft. These anomalies generally require rapid activation of the second turbine engine as they may be symptomatic of a significant loss of power from the first turbine engine.

[0058] Alternatively, the third condition C3 comprises the detection of flight conditions requiring emergency operation in nominal mode 10 of the second turbine engine, or comprises the reception by the second turbine engine of an activation instruction issued by an on-board computer and / or a pilot of the aircraft.

[0059] The rapid activation mode 20 is different from the two previous activation modes. In particular, the rapid activation mode 20 comprises the use of a specific starting system configured and sized to implement a fuel flow law allowing an activation time of between 5 and 15 seconds, around 10 seconds.

[0060] The rapid activation mode 20, however, has a significant impact on the aging of the second turboshaft engine and is only used in an emergency.

[0061] Among the three activation modes, the normal activation mode 16 is favored over the accelerated activation mode 18, and the accelerated activation mode 18 is also favored over the rapid activation mode 20. Indeed, an activation with low impact on aging is favored, although the first duration is longer than the second duration, and the second duration is longer than the third duration.

[0062] Optionally, the second turbine engine comprises an electric assistance machine and each of the normal activation, accelerated activation, or rapid activation modes comprises a step of assistance of the second turbine engine by the electric assistance machine of the aircraft.

[0063] In other words, the electric assistance machine participates in the rotational drive of the second turboshaft engine during its activation in order to accelerate said activation.

[0064] In a particular implementation mode where the standby mode of the second turbine engine implies that the combustion chamber is ignited, also called super idle mode, only the rapid activation mode requires the implementation of a step of assistance of the second turbine engine by the electric machine.

[0065] Advantageously, the flight parameters of the aircraft include the flight altitude of the aircraft, and the choice of the activation mode of the assistance engine, in other words of the second turbine engine, is made as a function of the possibility of maintaining the flight altitude of the aircraft for the time necessary for the activation of the second turbine engine.

[0066] Thus, the first, second and third conditions each comprise an additional condition of maintaining the flight altitude of the aircraft for the time necessary for the activation of the second turbine engine. Indeed, although the normal activation mode 16 is preferred because it has little impact on the aging of the second turbine engine, if this does not guarantee that the aircraft's flight altitude can be maintained, a faster activation mode will be preferred. The same reasoning applies to accelerated activation mode 18.

[0067] In addition, an activation mode may be interrupted in favor of another activation mode when the conditions of the other activation mode are met.

[0068] For example, when a normal activation mode 16 is in progress, but a significant failure is detected, a rapid activation mode 20 is implemented. However, if the specific starting system has a failure 22, an accelerated activation mode 18 is finally implemented.

[0069] In one embodiment, the method is implemented continuously by a computer of the second turbine engine as long as the second turbine engine is not in its nominal operating mode 10. Thus, the steps of comparing the parameters 2 and detecting failures 4 only stop when the second turbine engine is in its nominal operating mode 10.

Claims

Claims

1. Method for assisting propulsion by detecting a failure of a turbine engine of an aircraft operating in a nominal mode, the aircraft comprising a deactivated assistance engine, characterized in that it comprises a step (2) of comparing operating parameters of the turbine engine with the equivalent parameters of a model representative of a healthy turbine engine, a step (4) of detecting a failure of the turbine engine by detecting an anomaly of at least one operating parameter of the turbine engine, a step (6) of choosing an activation mode of the assistance engine as a function of the operating parameters of the turbine engine and / or flight parameters of the aircraft, and a step (8) of activating the assistance engine with the chosen activation mode,the operating parameters of the turbine engine or detectable failures comprising at least one of the following parameters: a deterioration in the quality of regulation of the performance of the turbine engine, an incompatibility between the nominal operating mode of the turbine engine and reduced margins in relation to the operating limits of the rotation speed, the internal temperature of the turbine engine, the torque supplied by the turbine engine, the acceleration or deceleration of said turbine engine, or the fuel flow limits, an abnormal vibration level of the turbine engine, an abnormally high temperature of an aircraft computer, a connection failure between an electronic control system of the turbine engine and a central computer of the aircraft, a detection of a fire in the turbine engine, a detection of a power oscillation of the turbine engine.,

2. Method according to claim 1, in which the flight parameters of the aircraft comprise the flight altitude of the aircraft, the choice of the mode of activation of the assistance motor being made according to the possibility of maintaining the flight altitude of the aircraft for the time necessary for the activation of the assistance motor.

3. Method according to one of claims 1 and 2, in which the step (2) of comparing operating parameters of the turbine engine with the equivalent parameters of a model representative of a healthy turbine engine is carried out by a computer of the turbine engine storing the model representative of a healthy turbine engine.

4. A method according to any one of claims 1 to 3, wherein step (4) of detecting a failure of the turbine engine comprises measuring a significant deviation between an operating parameter of the turbine engine and the equivalent parameter of the model representative of a healthy turbine engine.

5. Method according to any one of claims 1 to 4, in which the choice of the activation mode of the assistance motor is made between a normal activation mode (16) having a first activation duration, an accelerated activation mode (18) having a second activation duration and a rapid activation mode (20) having a third activation duration, the first duration being longer than the second duration, and the second duration being longer than the third duration, the normal activation mode (16) being preferred over the accelerated activation mode (18), the accelerated activation mode (18) also being preferred over the rapid activation mode (20).

6. A method according to any one of claims 1 to 5, wherein the assist engine is a second turbine engine.

7. Method according to claims 5 and 6, in which the normal activation mode (16) comprises a step of thermal stabilization at a predefined temperature of the assistance engine and is favored when no failure of the turbine engine is detected and an on-board computer and / or a pilot of the aircraft gives the instruction to the assistance engine.

8. Method according to any one of claims 1 and 7, in which the operating parameters of the turbine engine or the detectable failures further comprise the fuel flow rate, and / or the temperature and / or the rotational speed at the outlet of a high-pressure turbine of the turbine engine, and / or the pressure at the outlet of a compressor of the turbine engine, and / or the torque, and / or the flameout of a combustion chamber of the turbine engine, and / or a leak on an oil circuit, and / or a leak on a fuel circuit, and / or a detection of filings, and / or an inability to regulate the performance of the turbine engine, and / or an inconsistency between the power demand and the power supplied, and / or a potential impact on the trajectory of the aircraft.

9. A method according to any one of claims 5 to 8, wherein the assisting engine is a second turbine engine, and wherein the normal activation mode (16), the accelerated activation mode (18) and the rapid activation mode (20) comprise assisting the second turbine engine by an electrical machine of the aircraft, the mode normal activation mode (16) comprising thermal stabilization at a predefined temperature of the second turbine engine and the implementation of a fuel flow law allowing an activation time of between 1 and 3 minutes, the accelerated activation mode (18) comprising the implementation of a fuel flow law allowing an activation time of between 10 seconds and 1 minute, and the rapid activation mode (20) comprising the use of a specific starting system configured to implement an assistance torque and a fuel flow law allowing an activation time of the second turbine engine of between 5 seconds and 15 seconds.

10. Method according to any one of claims 1 to 9, in which the deactivated assistance motor corresponds to an operating mode of the assistance motor equivalent to a standby mode or to an intermediate mode between a standby mode and a nominal operating mode.

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