METHOD FOR CONTROLLING A GAS TURBINE ENGINE

Description

GENERAL TECHNICAL FIELD

[0001] The invention relates generally to the field of gas turbine engines, and more particularly to a method of controlling a gas turbine engine. STATE OF THE ART

[0002] In a known way, gas turbine engines can be used to provide thrust to an aircraft.

[0003] During its operation, a gas turbine engine heats up in a fairly uniform manner.

[0004] However, when the gas turbine engine is shut down, for example, after the aircraft has landed at an airport, the gas turbine engine cools unevenly. This is because some components cool down more quickly than others. If the gas turbine engine is left idle for a sufficiently long period (for example, a day), all of its components will gradually cool down until they reach the same temperature.

[0005] Conversely, if the gas turbine engine is restarted after being left idle for a short period (for example, one hour), some components, such as the high-pressure casing components, still exhibit significant temperature differences / heterogeneities upon restart. These temperature differences can lead to thermal expansion phenomena that may cause deformation of the engine components. For example, the high-pressure rotor line can flex between its bearings due to a vertical temperature gradient across these sections. This flexing of the rotor line causes, on the one hand, a reduction in static clearances between the moving blades and the casings, and on the other hand, creates an imbalance on the rotor / eccentric mass.If the gas turbine engine is restarted while the high-pressure drive shaft is bent, the engine's static thermal state, combined with a dynamic unbalance response, can lead to friction between the rotor blades of the high-pressure compressor or high-pressure turbine and the compressor or high-pressure turbine housing. If the rotor blades rub against the housing, this can damage the gas turbine engine and cause blade wear. Furthermore, if the housing is coated with an abradable material, this material will wear away, increasing the clearance between the rotor and stator of the gas turbine engine. This results in degraded engine performance.

[0006] Several solutions have already been considered to limit these temperature variations. Document EP3205847, for example, proposes using an electric motor connected to an air starter with a clutch to drive the high-pressure drive shaft via a gear train of an accessory gearbox (AGB) at very low speeds, below its idle speed: the motor is said to be in "rotisserie" mode. The slow rotation of the motor helps to homogenize the temperature of the engine components.

[0007] It is also known from document EP3415729 to control an electric motor via a FADEC (Full Authority Digital Engine Control) computer. The FADEC then controls the operation of both the gas turbine engine and the electric motor. When the FADEC is partially switched off, which corresponds to a stop of the aircraft, the electric motor rotates the drive shaft at reduced speed, i.e., in rotisserie mode.

[0008] Documents EP3205834 and EP3205858 propose using computer models to predict the required rotation time in a rotisserie for the electric motor connected to the air starter.

[0009] These known solutions aim to homogenize the temperature of the gas turbine engine components to prevent the drive shaft from bending by rotating it at low speed when the aircraft is stationary, i.e., when the gas turbine engine is off. However, these solutions require the aircraft to remain stationary for the time necessary for the temperature of the gas turbine engine components to homogenize, meaning that a certain period must be waited before the gas turbine engine can be restarted. Furthermore, the prior art solutions involve the use of dedicated systems. DESCRIPTION OF THE INVENTION

[0010] One aim of the invention is to remedy the aforementioned disadvantages by proposing a method for controlling a gas turbine engine that limits the waiting time required before restarting the gas turbine engine.

[0011] The present invention relates, in a first aspect, to a method for controlling a gas turbine engine, the gas turbine engine comprising a body including a compressor, a turbine and a drive shaft, the turbine being adapted to drive the compressor via the drive shaft, and an electric motor adapted to drive the drive shaft in rotation, the control method comprising the steps of: a) estimate a temperature gradient between a first part of the body and a second part of the body, b) compare the estimated temperature gradient to a predefined temperature gradient threshold, and c) if the temperature gradient is less than the predefined temperature gradient threshold, rotate the drive shaft with a rotational speed that varies over time according to a first rotational speed variation profile, d) if the temperature gradient is greater than the predefined temperature gradient threshold, control the electric motor to rotate the drive shaft in such a way as to vary the rotational speed of the drive shaft according to a second rotational speed variation profile over time, such that, when the rotational speed of the drive shaft is within a critical range of rotational speeds,an acceleration of the rotation of the drive shaft is greater than an acceleration of the rotation of the drive shaft according to the first profile of variation of rotational speed within the same critical interval of rotational speeds.

[0012] According to advantageous and non-limiting characteristics, taken alone or in any combination: The electric motor is a starter-generator.

[0013] The drive shaft is suitable for being driven in rotation around an axis of the gas turbine engine, and the estimated temperature gradient is obtained from temperature measurements in an upper part of the gas turbine engine body, located above the axis of the gas turbine engine, and from temperature measurements in a lower part of the gas turbine engine body, located below the axis of the gas turbine engine, when the gas turbine engine is attached to an aircraft.

[0014] The critical range of rotational speeds includes a rotational speed likely to generate resonance in the body, given the curvature of the drive shaft.

[0015] The critical rotational speed range is defined between a first rotational speed threshold and a second rotational speed threshold, the first rotational speed threshold and the second rotational speed threshold being predetermined and dependent on the gas turbine engine.

[0016] The power supplied by the electric motor to the drive shaft during step d) when the rotational speed of the drive shaft is within the critical range of rotational speeds is determined from the temperature gradient.

[0017] The control process includes a step of measuring a first value of a body vibration parameter, the power supplied by the electric motor being further determined from the first value of the body vibration parameter.

[0018] The piloting process includes a gas turbine engine start-up stage and a pre-start-stage stage in which the electric motor drives the drive shaft so as to rotate the drive shaft at a rotational speed of less than 10 revolutions per minute.

[0019] The piloting process includes a step of piloting the pressure of at least one damping fluid film disposed between the drive shaft and a housing of the gas turbine engine.

[0020] Step d) includes supplying power from the electric motor to the drive shaft when the rotational speed of the drive shaft is within the critical range, a step of measuring a second value of a body vibration parameter and, if the second value of the body vibration parameter is greater than a vibration parameter threshold, the power supplied by the electric motor to the drive shaft is decreased.

[0021] The body is a high-pressure body, the compressor is a high-pressure compressor, the turbine is a high-pressure turbine, and the drive shaft is a high-pressure drive shaft, and wherein the gas turbine engine further comprises a low-pressure body including a blower, a low-pressure turbine, and a low-pressure drive shaft, the low-pressure turbine being suitable for driving the blower via the low-pressure shaft.

[0022] According to a second aspect, the invention relates to a gas turbine engine, comprising a body including a compressor, a turbine and a drive shaft, the turbine being suitable for driving the compressor via the drive shaft, an electric motor suitable for driving the drive shaft in rotation, and a control module configured to control the gas turbine engine according to the steps of the process as previously defined.

[0023] According to advantageous and non-limiting features, taken alone or in any combination: The drive shaft is suitable for being driven in rotation about an axis of the gas turbine engine, and comprising at least two temperature sensors, including a first temperature sensor configured to measure a first temperature of an upper part of the body of the gas turbine engine, located above the axis of the gas turbine engine, and a second temperature sensor configured to measure a second temperature of a lower part of the body of the gas turbine engine, located below the axis of the gas turbine engine, when the gas turbine engine is attached to an aircraft.

[0024] According to a third aspect, the invention relates to an aircraft characterized in that it comprises the gas turbine engine described above. PRESENTATION OF THE FIGURES

[0025] Other features and advantages of the present invention will become apparent from the following description of a preferred embodiment. This description will be given with reference to the accompanying figures, including: There figure 1 schematically represents a gas turbine engine according to one possible embodiment of the invention; figure 2 schematically represents the position of the sensors in the gas turbine engine according to one possible embodiment of the invention; The figure 3 schematically represents the steps of a method for controlling a gas turbine engine according to one possible embodiment of the invention; The figure 4is a diagram schematically representing a rotational speed of a drive shaft of the gas turbine engine, an acceleration of the rotation of the drive shaft, as a function of time, according to two distinct control modes of the gas turbine engine and a power supply profile by the electric motor to the drive shaft. DETAILED DESCRIPTION Turbomachine

[0026] With reference to the Figure 1 The present invention relates to a gas turbine engine 1 which includes a blower 2, a low pressure body 3, a high pressure body 4, a combustion chamber 5 and a gas exhaust nozzle 6.

[0027] The high-pressure body 4 comprises a high-pressure compressor 41, a high-pressure turbine 42, and a high-pressure drive shaft 43 coupling the high-pressure turbine 42 to the high-pressure compressor 41. The low-pressure body 3 comprises a low-pressure compressor 31, a low-pressure turbine 32, and a low-pressure drive shaft 33 coupling the low-pressure turbine 32 to the low-pressure compressor 31, and extending inside the high-pressure drive shaft 43.

[0028] The high-pressure turbine 42 drives the high-pressure compressor 41 in rotation via the high-pressure drive shaft 43, while the low-pressure turbine 32 drives the low-pressure compressor 31 and the blower 2 in rotation via the low-pressure drive shaft 33.

[0029] The low-pressure drive shaft 33 is rotatably mounted about an axis of rotation X parallel to a longitudinal direction of the gas turbine engine 1. Similarly, the high-pressure drive shaft 43 is rotatably mounted about the axis of rotation X. The low-pressure drive shaft 33 and the high-pressure drive shaft 43 are coaxial. The high-pressure drive shaft 43 extends around the low-pressure drive shaft 33.

[0030] The high-pressure compressor 41 comprises a high-pressure compressor housing 412, a high-pressure compressor stator 414, fixedly mounted relative to the high-pressure compressor housing 412, and a high-pressure compressor rotor 416, adapted to be driven in rotation relative to the high-pressure compressor stator 414, about the axis of rotation X. Similarly, the low-pressure compressor 31 comprises a low-pressure compressor housing 312, a low-pressure compressor stator 314, fixedly mounted relative to the low-pressure compressor housing 312, and a low-pressure compressor rotor 316, adapted to be driven in rotation relative to the low-pressure compressor stator 314, about the axis of rotation X.

[0031] The high-pressure turbine 42 comprises a high-pressure turbine housing 422, a high-pressure turbine stator 424, fixedly mounted relative to the high-pressure turbine housing 422, and a high-pressure turbine rotor 426, adapted to be driven in rotation relative to the high-pressure turbine stator 424, about the axis of rotation X. Similarly, the low-pressure turbine 32 comprises a low-pressure turbine housing 322, a low-pressure turbine stator 324, fixedly mounted relative to the low-pressure turbine housing 322, and a low-pressure turbine rotor 326, adapted to be driven in rotation relative to the low-pressure turbine stator 324, about the axis of rotation X.

[0032] During operation, the fan 2 is driven in rotation by the low-pressure turbine 32, resulting in air circulation from the upstream to the downstream side of the gas turbine engine 1. A portion of the air passing through the fan 2 (primary flow) then passes successively through the low-pressure compressor 31, the high-pressure compressor 41, and is then injected into the combustion chamber 5. In the combustion chamber 5, the air is mixed with fuel. The combustion of the fuel generates exhaust gases that circulate successively through the high-pressure turbine 42, then through the low-pressure turbine 32, and are expelled via the exhaust nozzle 6.

[0033] Preferably, the gas turbine engine 1 includes bearings 7 that guide the rotation of the high-pressure drive shaft 43 and the low-pressure drive shaft 33. Typically, the gas turbine engine 1 includes two bearings 7 located downstream of the low-pressure compressor 31, one bearing 7 upstream of the high-pressure compressor 41, one bearing 7 downstream of the high-pressure turbine 42, and finally, one bearing 7 downstream of the low-pressure turbine 32. These bearings include high-pressure housings 72 adapted to guide the rotation of the high-pressure drive shaft 43, and low-pressure housings 74 adapted to guide the rotation of the low-pressure drive shaft 33. In general, the high-pressure housings 72 include the bearing 7 located upstream of the high-pressure compressor 41 and the bearing 7 located downstream of the high-pressure turbine 42.As for the low-pressure housing 7, they comprise the two housings 7 located downstream of the low-pressure compressor 31 and the housing 7 located downstream of the low-pressure turbine 32. The housings 7 are plain bearings or roller bearings. Advantageously, the housings 7 are interposed between a housing of the gas turbine engine 14 and the drive shaft 33, 43 which they support. In other words, each bearing 7 comprises an inner ring 73 and an outer ring 74, and one of the inner ring 73 and outer ring 74 is fixedly mounted on the gas turbine engine housing 14, while the other of the inner ring 73 and outer ring 74 is fixedly mounted on the drive shaft 33, 43. Preferably, the outer ring 74 is fixed to a bearing support integral with the gas turbine engine housing 14. Even more preferably, the bearing support extends from the outer ring 74 to a flange fixed relative to the gas turbine engine housing 14.The bearings 7 thus form interfaces between the drive shaft 33, 43 of the gas turbine engine 1 and the gas turbine engine housing 14. Preferably, a damping fluid film 75 is disposed between a bearing 7 and the gas turbine engine housing 14. Preferably, the damping fluid film 75 is a lubricating film such as an oil film. In a certain embodiment, the oil is supplied via an oil circuit common to the bearings 7.

[0034] The gas turbine engine 1 further comprises an electric motor 15. The electric motor 15 is capable of rotating one of the drive shafts 33, 43. In other words, the electric motor 15 allows one of the drive shafts 33, 43 to rotate relative to the housing of the gas turbine engine 14. Preferably, the electric motor 15 is capable of rotating the high-pressure drive shaft 43. The electric motor 15 can be directly connected to a drive shaft 33, 43 via a transmission gear or be connected to an accessory gearbox 150 (AGB).The accessory drive unit 150 generally comprises one or more gear trains which are suitable for being driven in rotation by a mechanical take-off by means of a right-angle gearbox on the drive shaft 33, 43 and a radial transmission shaft 151, on which various accessories such as high-pressure fuel pumps, lubrication pumps, etc. are coupled. The electric motor 15 can be connected to an accessory drive unit 150, itself connected to a drive shaft 33, 43 to be driven in rotation via a radial transmission shaft 151.

[0035] In one embodiment, the electric motor 15 is a starter-generator. A starter-generator is a motor, most often an electric motor, used to start the gas turbine engine 1. In one embodiment, the starter-generator 15 is an electric motor 15 capable of driving one of the drive shafts 33, 43 in rotation, for example in the case of a hybrid gas turbine engine 1. Hybrid gas turbine engines operate using both electrical energy and thermal energy from the combustion of gases in the combustion chamber 5. Preferably, the present invention does not involve adding a dedicated electric motor 15 to the gas turbine engine 1, but rather using an electric motor 15 already present in the gas turbine engine 1. This avoids adding bulk to the gas turbine engine 1, increasing its weight, and reducing its manufacturing cost.

[0036] Preferably, with reference to the Figure 2 The gas turbine engine 1 includes at least two temperature sensors 17. The temperature sensors 17 are configured to collect data relating to the temperature of one of the bodies 3, 4 of the gas turbine engine 1. Preferably, the temperature sensors 17 are attached to the high-pressure compressor housing 412 or to the high-pressure turbine housing 422.

[0037] Preferably, the two temperature sensors 17 are located on a housing surrounding the primary stream at a distance of less than 20 cm and, even more preferably, at a distance of less than 10 cm from the primary stream. The temperature sensors 17 are located opposite the high-pressure compressor 41 or the high-pressure turbine 42.

[0038] Preferably, the two temperature sensors 17 comprise a first temperature sensor 171 located on an upper part of the body 35, 45 of the gas turbine engine 1 and a second temperature sensor 172 located on a lower part of the body 36, 46 of the gas turbine engine 1.

[0039] More specifically, the first temperature sensor 171 is located on an upper part of the high-pressure compressor housing 412 or on an upper part of the high-pressure turbine housing 422, and the second temperature sensor 172 is located on a lower part of the high-pressure compressor housing 412 or the high-pressure turbine housing 422. These temperature sensors 17 allow for the collection of different temperature values ​​from the gas turbine engine 1, which will allow for the calculation of a temperature gradient between an upper part of the body 35, 45 of the gas turbine engine 1 and a lower part of the body 36, 46 of the gas turbine engine 1. The terms "upper" and "lower" are to be interpreted with regard to the gas turbine engine 1 being fixedly mounted on an aircraft 100, the aircraft 100 being placed on a horizontal ground.In other words, a lower portion of the body 36, 46 of the gas turbine engine 1 is located closer to the ground than an upper portion of the body 35, 45 of the gas turbine engine 1 when the gas turbine engine 1 is attached to the aircraft 100 and the aircraft 100 is resting on the horizontal ground. Preferably, the two temperature sensors 17 are located relative to each other in diametrically opposite positions with respect to the axis of rotation X.

[0040] In addition, advantageously, the gas turbine engine 1 includes a rotational speed sensor 18. Preferably, the rotational speed sensor 18 is configured to measure the rotational speed of the high-pressure drive shaft 43. The rotational speed sensor 18 is, for example, fixed to the high-pressure drive shaft 43.

[0041] Alternatively, the rotational speed sensor 18 can be fixed on the radial transmission shaft 151 connected to the accessory drive housing 150. In this case, the rotational speed of the high-pressure drive shaft 43 is obtained from the measurement of the rotational speed of the radial transmission shaft 151 and a gear ratio between the high-pressure drive shaft 43 and the radial transmission shaft 151.

[0042] Finally, preferably, the gas turbine engine 1 includes a vibration sensor. The vibration sensor 19 is configured to measure a vibration parameter of a body 3, 4 of the gas turbine engine 1. The vibration parameter is, for example, a displacement, a velocity, or an acceleration. This vibration parameter is expressed, for example, in mils DA when measuring a vibration parameter of a low-pressure body 3 and in inch / sec Peak when measuring a vibration parameter of a high-pressure body 4. The vibration sensor 19 is, for example, attached to a bearing support or to a flange of the low-pressure body 3 or the high-pressure body 4. The vibration sensor 19 is thus advantageously positioned to accurately capture / characterize the vibrations of the gas turbine engine 1. The position of the vibration sensor 19 can therefore vary depending on the model of the gas turbine engine 1. Process

[0043] With reference to the Figure 3The present invention relates to a method for controlling the gas turbine engine 1. This method is preferably implemented before the drive shaft 33, 43 of the gas turbine engine 1 reaches its idle speed. The idle speed of the drive shaft 33, 43 of the gas turbine engine 1 (hereinafter referred to as the "idle speed of the gas turbine engine 1" for brevity) is typically the minimum, steady-state speed of the engine. The idle speed of the gas turbine engine 1 is the rotational speed of the drive shaft 33, 43 of the gas turbine engine 1 when the aircraft 100 is on the ground, stationary, or moving, for example, during holding or taxiing phases.The idle speed of the gas turbine engine 1 is typically the minimum rotational speed at which the drive shaft 33, 43 of the gas turbine engine 1 rotates after the gas turbine engine 1 has started. The starting period of the gas turbine engine 1 is the time during which a starter-generator is required to rotate the drive shaft 33, 43. It is therefore understood that the idle speed of the gas turbine engine 1 is the speed reached by the gas turbine engine 1 when the starter-generator is switched off (i.e., the starting period is over).

[0044] The method first includes a step a) in which a temperature gradient is estimated between a first part of the body 35, 45 of the gas turbine engine 1 and the second part of the body 36, 46. Preferably, and more precisely, the temperature gradient is a temperature gradient between the first part of the high-pressure body 45 (and not the first part of the low-pressure body 35) of the gas turbine engine 1 and the second part of the high-pressure body 46 (and not the second part of the low-pressure body 36) of the gas turbine engine 1. Preferably, the temperature gradient is obtained from at least two temperature values ​​that are measured by the two temperature sensors 17 of the gas turbine engine 1 (i.e., the first temperature sensor 171 located on the first part of the body 35, 45 of the gas turbine engine 1 and the second temperature sensor 172 located on the second part of the body 36, 46 of the gas turbine engine 1).Preferably, the first part of the high-pressure body 45 of the gas turbine engine 1 is an upper part of the high-pressure body 45 of the gas turbine engine 1 and the second part of the high-pressure body 46 of the gas turbine engine 1 is a lower part of the high-pressure body 46 of the gas turbine engine 1.

[0045] Thus, preferably, the estimated temperature gradient is obtained from temperature values ​​measured in the upper part of the body 35, 45 of the gas turbine engine 1, located above the axis of the gas turbine engine 1, and from temperature values ​​measured in the lower part of the body 36, 46 of the gas turbine engine 1, located below the axis of the gas turbine engine 1, when the gas turbine engine 1 is attached to an aircraft 100. The temperature gradient is thus obtained from at least two temperature values, including a first temperature value from an upper part of the body 35, 45 of the gas turbine engine 1 and a second temperature value from a lower part of the body 36, 46 of the gas turbine engine 1.Preferably, the estimated temperature gradient is calculated from at least two temperature values: a first temperature value for an upper portion of the high-pressure body 45 (preferably an upper portion of the high-pressure compressor 415) and a second temperature value for a lower portion of the high-pressure body 46 (preferably a lower portion of the high-pressure compressor 417). Thus, the estimated temperature gradient advantageously represents the temperature difference between the upper and lower portions of the high-pressure body 46. Following the shutdown of the gas turbine engine 1, the temperatures of the bodies 3 and 4 of the gas turbine engine 1 become heterogeneous. In particular, the temperature of the high-pressure body 4 becomes heterogeneous. Indeed, during operation of the gas turbine engine 1, the high-pressure body 4 is significantly hotter than the low-pressure body 3.During the cooling of the gas turbine engine 1, the lower part of the high-pressure body 46 cools down faster than the upper part of the high-pressure body 45. This is simply explained by the fact that the air in the stationary gas turbine engine stratifies, with the lighter, warmer air rising. Thus, the air heated by the cooling components of the gas turbine engine 1 tends to rise. Consequently, the upper part of the high-pressure body 45 is heated by hot air, while the lower part of the high-pressure body 46 is less subject to this phenomenon and therefore cools down faster than the upper part of the high-pressure body 45. After approximately one to two hours of the gas turbine engine 1 being off, the temperature difference between a point on the upper part of the high-pressure body 45 and a point on the lower part of the high-pressure body 46 is typically between 30 and 50 degrees.The estimated temperature gradient of the high-pressure body 4 is therefore representative of the thermal heterogeneity of the high-pressure body 4 due to this phenomenon.

[0046] Preferably, step a) of the process is carried out before or during the starting of the gas turbine engine 1. Typically, the starting of the gas turbine engine 1 includes, but is not limited to, the following steps: The drive shaft 33, 43 of the gas turbine engine 1 is rotated, preferably by the starter-generator, to a starting speed. Ambient air is compressed by the rotating compressors 31, 41 and the compressed air enters the combustion chamber 5 into which fuel is injected. The compressed air-fuel mixture is combusted, generating highly expanded gas flows. The gas flows pass through the turbines 32, 42 and then exit the gas turbine engine 1 via the exhaust nozzle 6, and the drive shaft 33, 43 of the gas turbine engine 1 is rotated by these gas flows. The starter-generator is then switched off. The start-up is complete when the gas turbine engine 1 is operating independently, that is, when the drive shaft 33, 43 of the gas turbine engine 1 is rotating without requiring a drive from the starter-generator. Once the start-up is complete, the gas turbine engine 1 is considered to be running.

[0047] The temperature gradient can therefore be estimated even before the gas turbine engine 1 starts, i.e., before the drive shaft 33, 43 of the gas turbine engine 1 begins to rotate to a starting speed. The temperature gradient can also be estimated during the start-up of the gas turbine engine 1. By estimating the temperature gradient before or during the start-up of the gas turbine engine 1, the estimated temperature gradient is representative of the temperature heterogeneity of the body 3, 4 of the gas turbine engine 1 during or after a shutdown of the gas turbine engine 1.

[0048] It should be noted that the thermal gradient estimation can be performed using different types of temperature measurements. For example, in one embodiment, the measurements can be taken on a test bench, i.e., when the gas turbine engine 1 is not mounted on an aircraft and is being tested. In this case, the temperature measurements for estimating the temperature gradient can be taken after the gas turbine engine 1 has already started to determine the evolution of the temperature heterogeneity of the body 3, 4 of the gas turbine engine 1 during startup, or after startup and after shutdown of the gas turbine engine 1. In a second embodiment, the temperature measurements are taken in a fleet, i.e., when the gas turbine engine 1 is mounted on an aircraft and the aircraft is in operation. The temperature measurements are therefore taken under "real-world conditions," i.e.under traditional operating conditions of a gas turbine engine 1. In this case, temperature values ​​are recorded before and during the initial moments of rotation of the drive shaft 33, 43 of the gas turbine engine 1 up to a time t (less than 5 min, preferably less than 1 min, and even more preferably less than 15 seconds). Based on the gradient estimated from the measurements, the variation profile is selected according to step b). Time t is a function of the engine speed and the mode positions. According to a third embodiment, the temperature measurements are taken on a digital twin of the gas turbine engine 1, thus in a simulated manner. In other words, the present method can be implemented on a gas turbine engine 1 under various conditions (test bench, fleet, simulation, etc.) which are not limited to the present description.

[0049] Then, in step b), the estimated temperature gradient is compared to a predefined temperature gradient threshold. The predefined temperature gradient threshold corresponds to a temperature gradient of a body 3, 4 of the gas turbine engine 1 beyond which a drive shaft 33, 43 of the gas turbine engine 1 is bent such that, when the drive shaft 33, 43 is rotated, damage to the gas turbine engine 1 occurs. Here, the drive shaft 33, 43 in question is preferably the high-pressure drive shaft 43. More precisely, the predefined temperature gradient threshold corresponds to a temperature gradient of the body 3 or 4 of the gas turbine engine 1 beyond which the drive shaft 33 or 43 is bent such that, when rotated, it vibrates to the point of causing damage to the gas turbine engine 1.In other words, the temperature gradient threshold corresponds to a temperature gradient of the body 3, 4 of the gas turbine engine 1 beyond which the drive shaft 33, 43 is bent in such a way that, when driven in rotation, it is likely to generate resonance of the body 3 or 4, which leads to damage to the gas turbine engine 1. Thus, the temperature gradient threshold is associated with a bending of a drive shaft 33, 43, a bending beyond which it is estimated that the drive shaft 33, 43 will be bent to the point of causing damage to the gas turbine engine 1. Consequently, the estimated temperature gradient allows us to determine an estimate of the bending of the drive shaft 33, 43, for example, a radius of curvature. More specifically, the estimated temperature gradient allows us to determine a deflection, i.e. a maximum displacement value, of the drive shaft 33, 43.Then, the deflection of the drive shaft 33, 43 is used to determine the deflection. The temperature gradient then indicates the degree of deflection of the drive shaft 33, 43. Preferably, a model is used to determine the deflection of the drive shaft 33, 43 based on the estimated temperature gradient. By comparing the estimated temperature gradient to the temperature gradient threshold, it is possible to determine whether the drive shaft 33, 43 is deflected to such an extent that damage to the gas turbine engine 1 will occur.

[0050] If the estimated temperature gradient is below the predefined temperature gradient threshold, in step c), the drive shaft 33, 43 is rotated at a speed that varies over time according to a first speed variation profile. This first speed variation profile is a conventional speed variation profile for reaching the idle speed of the drive shaft 33, 43. In fact, if the temperature gradient is below the predefined temperature gradient threshold, it is assumed that the drive shaft 33, 43 is not curved to the point of potentially causing resonance of the body 3, 4 because the body 3, 4 has cooled sufficiently and is cooled uniformly enough. Consequently, the gas turbine engine 1 is driven to operate normally, that is, in a conventional manner for the operation of a gas turbine engine 1.As explained previously, the present method is implemented before the gas turbine engine 1 operates at idle speed, that is, before the drive shaft 33, 43 rotates at idle speed. The first profile of the rotational speed variation over time is a profile of the rotational speed variation of a conventional drive shaft 33, 43, before reaching idle speed. Typically, the drive shaft 33, 43 is driven to rotate with a rotational speed that increases over time due to an acceleration applied to the drive shaft 33, 43 by the starter-generator. Generally, acceleration is applied to the drive shaft 33, 43 by the starter-generator during starting to enable the drive shaft 33, 43 to accelerate until it reaches an idle rotation speed.An example of a first rotational speed variation profile C1 is illustrated in . Figure 4 Curves C1 and C2 of the Figure 4 illustrate changes in rotational speed over time.

[0051] On the other hand, if the estimated temperature gradient is greater than the predefined temperature gradient threshold, in a step d), the starter-generator 15 is controlled to drive the drive shaft 33, 43 in such a way as to vary the rotational speed of the drive shaft 33, 43 according to a second profile of variation of rotational speed over time such that, when the rotational speed of the drive shaft 33, 43 is within a critical interval of rotational speeds, an acceleration of the rotation of the drive shaft 33, 43 is greater than an acceleration of the rotation of the drive shaft 33, 43 according to the first profile of variation of rotational speed in the same critical interval of rotational speeds.Indeed, damage to the gas turbine engine 1 occurs when the rotational speed of the curved drive shaft 33, 43 falls within a critical range of rotational speeds because the drive shaft 33, 43, and therefore the body 3, 4, vibrates more intensely. More specifically, if the curved drive shaft 33, 43 rotates at a speed within the critical range of rotational speeds, the body 3, 4 is likely to resonate. By implementing step d), the drive shaft 33, 43 rotates for the shortest possible time at a speed within the critical range of rotational speeds, thereby limiting damage to the gas turbine engine 1.In other words, the shorter the period during which the rotational speed of the drive shaft 33, 43 is within the critical range of rotational speeds, the shorter the period during which the drive shaft 33, 43 vibrates intensely, and the shorter and less intense the period during which the drive shaft 33, 43 generates damage to the gas turbine engine 1. The period during which the drive shaft 33, 43 rotates at a speed within the critical range of rotational speeds is reduced compared to the conventional drive profile of the drive shaft 33, 43. Consequently, the adverse consequences (i.e., damage to the gas turbine engine 1) of the curved drive shaft 33, 43 will be reduced.Thus, step d) consists of controlling the electric motor 15 so that, when the rotational speed of the drive shaft 33, 43 is within the critical range of rotational speeds, it accelerates the drive shaft 33, 43 if the temperature gradient exceeds the temperature gradient threshold. Therefore, the rotational speed of the drive shaft 33, 43 varies according to a second profile of rotational speed variation over time, which differs from the first profile of rotational speed variation over time. Indeed, if the rotational speed of the drive shaft 33, 43 varies according to the second profile of rotational speed variation, then the speed of the drive shaft 33, 43, when it is within the critical range of rotational speeds, increases more rapidly than if the rotational speed of the drive shaft 33, 43 varied according to the first profile.An example of a second rotational speed variation profile C2 is illustrated in . Figure 4 If the temperature gradient exceeds the temperature gradient threshold, the drive shaft 33, 43 is accelerated, when its rotational speed is within the critical range of rotational speeds, by the electric motor 15, and this acceleration is greater than the acceleration of the drive shaft 33, 43 if the temperature gradient were below the temperature gradient threshold. To apply this acceleration, the electric motor 15 applies power to the drive shaft 33, 43. Curve C3 of the Figure 4illustrates this power application. Curve C3 is a schematic binary curve (state A when no power is applied and state B when power is applied) which schematically represents the supply of power to the drive shaft 33, 43 when the rotational speed of the drive shaft 33, 43 reaches the lower bound of the critical rotational speed range.

[0052] The present method does not require the aircraft 100 to wait on the tarmac while the temperature of body 3, 4 homogenizes before the gas turbine engine 1 of the aircraft 100 can be started. Step d) of the present method is implemented during the starting process of the gas turbine engine 1. The control method just described therefore makes it possible to limit the damage to the gas turbine engine 1 caused by the thermal heterogeneity of body 3, 4 without preventing the starting of the gas turbine engine 1.

[0053] In step d), the electric motor 15 accelerates the drive shaft 33, 43 to more quickly pass through the critical rotational speed range. The critical rotational speed range is the set of rotational speeds between a first rotational speed threshold and a second rotational speed threshold. The first rotational speed threshold is the lower bound of the critical rotational speed range. The second rotational speed threshold is the upper bound of the critical rotational speed range. As soon as the rotational speed of the drive shaft 33, 43 reaches the first rotational speed threshold, the electric motor 15 accelerates the drive shaft 33, 43 so that the drive shaft 33, 43 reaches a rotational speed higher than the second rotational speed threshold.For example, the electric motor 15 can accelerate the drive shaft 33, 43 when it is rotating at a speed of 3500 revolutions per minute until the drive shaft 33, 43 reaches a speed of 7000 revolutions per minute. The rotational speed sensor 18 detects when the drive shaft 33, 43 is rotating at a speed equal to the first or second rotational speed threshold. Therefore, if the estimated temperature gradient is greater than the predefined temperature gradient threshold, the electric motor 15 accelerates the drive shaft 33, 43 as soon as it reaches a rotational speed equal to the first rotational speed threshold.Thus, between the first rotation speed threshold and the second rotation speed threshold, the rotation speed increases faster according to the second rotation speed variation profile than according to the first rotation speed variation profile.

[0054] The first and second rotational speed thresholds are preferably predetermined and depend on the gas turbine engine 1. The critical range of rotational speeds depends on the type (i.e., the model) of the gas turbine engine 1. The rotational speed likely to generate a resonant input of the body (3, 4), which would cause degradation of the gas turbine engine 1, depends on the geometry of the gas turbine engine 1 and therefore on the type of gas turbine engine 1. Thus, the first and second rotational speed thresholds are predetermined for a gas turbine engine 1.

[0055] The electric motor 15 is configured to supply power to the drive shaft 33, 43, thereby accelerating the rotation of the drive shaft 33, 43. The power applied by the electric motor 15 to the drive shaft 33, 43 directly results in the acceleration of the drive shaft 33, 43 caused by the electric motor 15. More precisely, the greater the power applied by the electric motor 15 to the drive shaft 33, 43, the greater the acceleration of the drive shaft 33, 43. The power applied by the electric motor 15 to the drive shaft 33, 43 depends primarily on the specific electric motor 15 used. Indeed, the power applied by the electric motor 15 to the drive shaft 33, 43 cannot exceed the maximum power applicable by the electric motor 15.In one embodiment, the power applied by the electric motor 15 to the drive shaft 33, 43 is fixed. For example, the electric motor 15 applies a constant power of 350kW to the drive shaft 33, 43.

[0056] Preferably, the power supplied by the electric motor 15 to the drive shaft 33, 43 during step d) when the rotational speed of the drive shaft 33, 43 is within the critical range of rotational speeds is determined from the temperature gradient. In this way, the power supplied by the electric motor 15 to the drive shaft 33, 43 depends on the deflection of the drive shaft 33, 43 caused by the temperature gradient. The greater the temperature gradient, the greater the deflection of the drive shaft 33, 43 and the greater the vibrations of the body 3, 4 are likely to be when the drive shaft 33, 43 rotates at a speed within the critical range of rotational speeds. Consequently, the greater the temperature gradient, the greater the potential for damage to the gas turbine engine 1.Therefore, it is all the more desirable that the drive shaft 33, 43 rotate for the shortest possible time at a rotational speed within the critical range. Thus, if the temperature gradient is significant, it is desirable to increase the rotational speed of the drive shaft 33, 43 more sharply when it is rotating at a speed within the critical range. If the temperature gradient is small, the damage to the gas turbine engine 1 is less significant, and it is not as necessary for the drive shaft 33, 43 to rotate for the shortest possible time at a speed within the critical range.In short, the greater the temperature gradient, the greater the power applied by the electric motor 15 to the drive shaft 33, 43 is to accelerate the rotation of the drive shaft 33, 43 in the critical range of rotational speeds.

[0057] In a preferred embodiment, the method includes a step of measuring a first value of a vibration parameter of body 3, 4, and the power supplied by the electric motor (15) is further determined from this first value of the vibration parameter of body (3, 4). In other words, the power applied by the electric motor 15 to the drive shaft 33, 43 is determined based on at least one measurement of at least one vibration parameter of a body 3, 4 of the gas turbine engine 1. In fact, when the gas turbine engine 1 starts, the value of the vibration parameter allows the intensity of the vibrations of body 3, 4 to be estimated when the drive shaft 33, 43 rotates at a speed within the critical range of rotational speeds. Thus, the power supplied by the electric motor 15 to the drive shaft 33, 43 is adjusted to this estimate of the vibration intensity.The reasoning here is the same as in the previous paragraph concerning the determination of the power supplied by the electric motor 15 to the drive shaft 33, 43 based on the temperature gradient. If the vibrations are expected to be particularly intense, it is desirable to increase the rotational speed of the drive shaft 33, 43 more sharply when the drive shaft 33, 43 is rotating at a speed within the critical range of rotational speeds. Thus, if the vibrations are expected to be particularly intense, the power supplied by the electric motor 15 to the drive shaft 33, 43 must be greater than if the vibrations were expected to be weak, because the drive shaft 33, 43 must be accelerated more when passing through the critical range of rotational speeds.

[0058] Preferably, the power supplied by the electric motor 15 to the drive shaft 33, 43 is adjusted to avoid causing additional vibrations that would damage the gas turbine engine 1. Indeed, the acceleration of the drive shaft 33, 43 controlled by the electric motor 15 can result in vibrations of the body 3, 4 and thus damage to the gas turbine engine 1. It is therefore desirable that the intensity of the vibrations due to the acceleration of the drive shaft 33, 43 does not exceed a vibration parameter threshold, and therefore that the acceleration of the drive shaft 33, 43 must not exceed a certain corresponding acceleration threshold. Consequently, during the implementation of step d), the second values ​​of a vibration parameter of the body 3, 4 are measured by a vibration sensor 18 and then compared to the vibration parameter threshold.In this way, it is verified that the intensity of the body's vibrations does not exceed the vibration parameter threshold. If a second value of a body vibration parameter 3, 4 is greater than the vibration parameter threshold, the electric motor 15 is controlled so that the power supplied to the drive shaft 33, 43 is reduced and, consequently, the rotational acceleration of the drive shaft 33, 43 is reduced.

[0059] In a preferred embodiment, the method for controlling the gas turbine engine 1 includes a step a0), prior to step a), in which the electric motor 15 drives the drive shaft 33, 43 so as to rotate the drive shaft 33, 43 at a low rotational speed, less than 10 revolutions per minute. Preferably, the reduced rotational speed is less than 5 revolutions per minute. Even more preferably, the reduced rotational speed is less than 2 revolutions per minute. Preferably, step a0) is carried out before the gas turbine engine starts, i.e., when the gas turbine engine is stopped. Step a0) is, for example, carried out when the aircraft 100 is stationary on the tarmac. Step a0) allows us to limit thermal heterogenization, i.e. reduce the temperature gradient, of body 3, 4 and therefore of drive shaft 33, 43.Indeed, by continuing to rotate even when the gas turbine engine 1 is stopped, the body 3, 4, and therefore the drive shaft 33, 43, cools more evenly than if it were not rotating. This technique is called "rotisserie." By combining the rotisserie effect and the acceleration of the drive shaft 33, 43 implemented in step d), the detrimental effects of the bent drive shaft 33, 43 on the gas turbine engine 1 are significantly reduced. The rotisserie limits the deflection of the drive shaft 33, 43, and the acceleration of the drive shaft 33, 43 reduces the time during which the bent drive shaft 33, 43 damages the gas turbine engine 1.

[0060] Preferably, the process further includes a step of controlling the pressure of at least one damping fluid film 75 located between the drive shaft 33, 43 and the housing of the gas turbine engine 14. More precisely, these pressure-controlled damping fluid films are preferably located between the high-pressure housing bearings 4 and the high-pressure housing 4. Advantageously, these damping fluid films are lubricating layers such as oil. The purpose of these damping fluid films is to dampen the vibrations of the drive shaft 33, 43. Indeed, when the drive shaft 33, 43 vibrates, for example when it is bent, all the components mounted on the drive shaft 33, 43 vibrate, and this causes damage to the gas turbine engine 1. It is therefore desirable to dampen these vibrations. Fluid damping films between the bearings 7 and the housing of the gas turbine engine 14 are provided for this purpose.However, the pressure of the damping fluid films is generally variable because it depends on the rotational speed of the drive shaft 33, 43. For example, the pressure of the damping fluid films can vary between two and eight bar. Thus, in the prior art, the damping fluid films do not always optimally dampen the vibrations of the drive shaft 33, 43. Therefore, advantageously, the pressure of the damping fluid films is controlled. The pressure can, for example, be controlled by an electric pump 20. When the damping fluid film 75 consists of a damping fluid, such as oil, the electric pump 20 is connected to the damping fluid circuit to control its pressure. According to a particular embodiment, the electric pump 20 is connected to the damping fluid circuit that supplies damping fluid to the damping fluid films of the high-pressure housing bearings 4.The damping fluid circuit for the damping fluid films of the high-pressure bearing bodies 72 and the low-pressure bearing bodies 74 split at a certain point to create a separate damping fluid circuit for the high-pressure bearing bodies 72 and a separate damping fluid circuit for the low-pressure bearing bodies 74. In this way, the electric pump 20 is connected only to the damping fluid circuit for the high-pressure bearing bodies 72 to control the pressure of the damping fluid films in the high-pressure bearing bodies 72. Controlling the pressure means regulating the pressure. By controlling the pressure of the damping fluid films, the vibration damping of the drive shaft 33, 43 is adjusted to the intensity of the vibrations of the drive shaft 33, 43.For example, if the vibrations intensify, it may be necessary to increase the pressure of the damping fluid film 75 so that it can absorb these vibrations as much as possible. Thus, controlling the pressure helps to limit the vibrations of the drive shaft 33, 43, which can be caused by the fact that the drive shaft 33, 43 is curved. Consequently, the damage caused by these vibrations is limited.

[0061] This description refers to the body 3, 4 and the drive shaft 33, 43. Preferably, the body 3, 4 in question is the high-pressure body 4, and the drive shaft 33, 43 is therefore the high-pressure drive shaft 43. Indeed, as explained previously, thermal heterogeneization during the cooling of the gas turbine engine is greater for the high-pressure body 4 than for the low-pressure body 3. It is therefore the high-pressure drive shaft 43 that is more likely to be bent to the point that the high-pressure body 4 is susceptible to resonance. Thus, the present method applies primarily to the high-pressure body 4.

[0062] However, the invention is not limited to the embodiment described and shown in the accompanying figures. Modifications remain possible, particularly with regard to the constitution of the various technical features or by substitution of technical equivalents, without departing from the scope of protection of the invention as defined by the claims.

Claims

1. Method for controlling a gas turbine engine (1), the gas turbine engine (1) comprising a body (3, 4) comprising a compressor (31, 41), a turbine (32, 42), and a drive shaft (33, 43), the turbine (32, 42) being adapted to drive the compressor (31, 41) via the drive shaft (33, 43), and an electric motor (15) adapted to drive the drive shaft (33, 43) in rotation, the control method comprising the steps of: a) estimating a temperature gradient between a first part of the body (35, 45) and a second part of the body (36, 46), b) comparing the estimated temperature gradient to a predefined temperature gradient threshold, and c) if the temperature gradient is less than the predefined temperature gradient threshold, driving the drive shaft (33, 43) to rotate at a rotational speed that varies over time according to a first rotational speed variation profile, d) if the temperature gradient is greater than the predefined temperature gradient threshold, controlling the electric motor (15) to rotate the drive shaft (33, 43) so as to vary the rotational speed of the drive shaft (33, 43) according to a second rotation speed variation profile over time, such that, when the rotation speed of the drive shaft (33, 43) is within a critical range of rotation speeds, an acceleration of the rotation of the drive shaft (33, 43) is greater than an acceleration of the rotation of the drive shaft (33, 43) according to the first rotation speed variation profile in the same critical range of rotation speeds.

2. Control method according to claim 1, wherein the electric motor (15) is a starter-generator.

3. Control method according to one of claims 1 and 2, wherein the drive shaft (33, 43) is capable of being driven in rotation about an axis of the gas turbine engine (1), and the estimated temperature gradient is obtained from temperature measurements in an upper part of the body (35, 45) of the gas turbine engine (1), located above the axis of the gas turbine engine (1), and temperature measurements in a lower part of the body (36, 46) of the gas turbine engine (1), located below the axis of the gas turbine engine (1), when the gas turbine engine (1) is attached to an aircraft (100).

4. A control method according to any one of claims 1 to 3, wherein the critical range of rotational speeds includes a rotational speed capable of causing the body (3, 4) to enter resonance, taking into account a curvature of the drive shaft (33, 43).

5. A control method according to any one of claims 1 to 4, wherein the critical rotation speed range is defined between a first rotation speed threshold and a second rotation speed threshold, the first rotation speed threshold and the second rotation speed threshold being predetermined and dependent on the gas turbine engine (1).

6. Control method according to any of claims 1 to 5, wherein a power supplied by the electric motor (15) to the drive shaft (33, 43) in step d) when the rotational speed of the drive shaft (33, 43) is within the critical rotational speed range is determined from the temperature gradient.

7. Control method according to claim 6, comprising a step of measuring a first value of a vibration parameter of the body (3, 4), the power supplied by the electric motor (15) being further determined from the first value of the vibration parameter of the body (3, 4).

8. A control method according to any one of claims 1 to 7, comprising a step of starting the gas turbine engine (1) and a step prior to the step of starting the gas turbine engine (1) in which the electric motor (15) rotates the drive shaft (33, 43) so as to rotate the drive shaft (33, 43) at a rotational speed of less than 10 revolutions per minute.

9. A control method according to any one of claims 1 to 8, comprising a step of controlling the pressure of at least one damping fluid film (75) disposed between the drive shaft (33, 43) and a casing of the gas turbine engine (1).

10. A control method according to any one of claims 1 to 9, wherein step d) comprises supplying power from the electric motor (15) to the drive shaft (33, 43) when the rotational speed of the drive shaft (33, 43) is within the critical range, a step of measuring a second value of a vibration parameter of the body (3, 4) and, if the second value of the vibration parameter of the body (3, 4) is greater than a vibration parameter threshold, the power supplied by the electric motor (15) to the drive shaft (33, 43) is decreased.

11. Control method according to one of claims 1 to 10, wherein the body (3, 4) is a high-pressure body (4), the compressor (31, 41) is a high-pressure compressor (41), the turbine (32, 42) is a high-pressure turbine (42), and the drive shaft (33, 43) is a high-pressure drive shaft (43), and wherein the gas turbine engine (1) further comprises a low-pressure body (3) comprising a fan (2), a low-pressure turbine (32) and a low-pressure drive shaft (33), the low-pressure turbine (32) being capable of driving the fan (2) via the low-pressure shaft (33).

12. Gas turbine engine (1), comprising a body (3, 4) comprising a compressor (31, 41), a turbine (32, 42) and a drive shaft (33, 43), the turbine (32, 42) being adapted to drive the compressor (31, 41) via the drive shaft (33, 43), an electric motor (15) capable of driving the drive shaft (33, 43) in rotation, and a control module configured to control the gas turbine engine (1) according to the steps of the method defined by one of claims 1 to 11.

13. Gas turbine engine (1) according to claim 12, wherein the drive shaft (33, 43) is adapted to be driven in rotation about an axis of the gas turbine engine (1), and comprising at least two temperature sensors (17, 171, 172), including a first temperature sensor (171) configured to measure a first temperature of an upper part of the body (35, 45) of the gas turbine engine (1), located above the axis of the gas turbine engine (1), and a second temperature sensor (172) configured to measure a second temperature of a lower part of the body (36, 46) of the gas turbine engine (1), located below the axis of the gas turbine engine (1), when the gas turbine engine (1) is attached to an aircraft (100).

14. Aircraft (100) characterized in that it comprises the gas turbine engine (1) according to one of claims 12 and 13.