Variable-pitch stator vane control

Abstract

The present disclosure relates to a computer-implemented method comprising: determining a value of a target pitch angle of a vane (40b, 50b) of a stator portion of a stage of a compressor of an aeronautical propulsion system of an aircraft, based on a value of a parameter (P) representative of wear of the aeronautical propulsion system; and generating a control signal (S) configured to control a control device (400) controlling the vane (40b, 50b) such that the control device (400) adjusts a pitch angle (α) of the vane (40b, 50b) to the determined value of the target pitch angle.

Description

[0001] Variable pitch stator blade control

[0002] TECHNICAL FIELD

[0003] This presentation concerns the field of aeronautics. More specifically, this presentation concerns the control of variable-pitch stator blades in an aircraft engine.

[0004] STATE OF THE ART

[0005] Climate change is a major concern for many legislative and regulatory bodies worldwide. Indeed, various restrictions on carbon emissions have been, are being, or will be adopted by different countries. In particular, an ambitious standard applies to both new types of aircraft and those already in operation, requiring the implementation of technological solutions to bring them into compliance with current regulations. Civil aviation has been actively contributing to the fight against climate change for several years now.

[0006] Technological research efforts have already led to very significant improvements in the environmental performance of aircraft. The Applicant takes into account the factors impacting all phases of design and development in order to obtain less energy-intensive and more environmentally friendly aeronautical components and products, whose integration and use in civil aviation have moderate environmental consequences, with the aim of improving aircraft energy efficiency.

[0007] Consequently, the Applicant is constantly working to reduce its negative climate impact by using methods and operating virtuous development and manufacturing processes that minimize greenhouse gas emissions to the minimum possible in order to reduce the environmental footprint of its activity.

[0008] This sustained research and development work focuses on new generations of aeronautical engines, the weight reduction of aircraft, particularly through the materials used and lighter on-board equipment, the development of the use of electrical technologies to provide propulsion, and, as essential complements to technological progress, aviation biofuels.

[0009] To guard against compressor surge, aircraft engines are generally designed with an operating margin, so that, in their most degraded state, their operating line is positioned some distance from the surge line. This operating margin, however, reduces their efficiency when new.

[0010] SUMMARY OF THE PRESENTATION

[0011] One aim of this presentation is to improve the efficiency of an aircraft engine without increasing the risk of compressor surge.

[0012] To this end, a computer-implemented process is proposed, according to one aspect of this presentation, comprising:

[0013] the determination of a target pitch angle value for a stator blade of a compressor stage in an aircraft propulsion system, based on a value of a parameter representative of wear in the aircraft propulsion system; and

[0014] the generation of a control signal configured to control a blade piloting device so that the piloting device adjusts a blade pitch angle to the determined value of the target pitch angle.

[0015] The determination of the target pitch angle value can also be implemented from an altitude value of the aeronautical propulsion system.

[0016] The determination of the target pitch angle value can also be implemented from a Mach number value of the aircraft.

[0017] The determination of the target pitch angle value can also be implemented from a value of an aeronautical propulsion system regime.

[0018] Determining the target pitch angle value may include determining a control law that provides target pitch angle values ​​from values ​​of an aeronautical propulsion system regime;

[0019] in which the control law is determined from the value of the parameter representing the wear of the aeronautical propulsion system, a value of an altitude of the aeronautical propulsion system, and a value of a Mach number of the aircraft; in which the value of the target pitch angle is determined from the control law and a value of the regime of the aeronautical propulsion system.

[0020] The value of the parameter representing the wear of the aeronautical propulsion system can be a number of operating cycles of the aeronautical propulsion system since its commissioning. The determination of the target pitch angle value may include determining a value for the exhaust gas temperature of the aeronautical propulsion system at the time of its commissioning;

[0021] in which the value of the exhaust gas temperature of the aeronautical propulsion system when it is put into service is determined from a value of a regime of the aeronautical propulsion system, and from a reference law providing values ​​of the exhaust gas temperature of the aeronautical propulsion system when it is put into service as a function of values ​​of the regime of the aeronautical propulsion system;

[0022] in which the value of the parameter representing the wear of the aeronautical propulsion system is a difference between a value of an exhaust gas temperature of the aeronautical propulsion system and the determined value of the exhaust gas temperature of the aeronautical propulsion system at the time of its commissioning.

[0023] According to another aspect of this presentation, a computer program product is proposed comprising instructions which, when the program is executed by a computer, cause the computer to implement the process according to this presentation.

[0024] According to another aspect of this presentation, a data processing device is proposed for an aeronautical propulsion system comprising a processor configured to implement the process according to this presentation.

[0025] According to another aspect of this presentation, an aeronautical propulsion system is proposed, having a longitudinal axis along which the aeronautical propulsion system extends, the aeronautical propulsion system comprising:

[0026] the data processing system as described in this presentation;

[0027] a compressor comprising a stage, the stage comprising a stator portion, the stator portion comprising a blade having a pitch axis along which the blade extends, the pitch axis extending radially with respect to the longitudinal axis, the blade being configured to pivot about the pitch axis so as to present a pitch angle; and

[0028] a control device configured to adjust the blade pitch angle to the target pitch angle value from the control signal generated by the data processing device.

[0029] DESCRIPTION OF THE FIGURES

[0030] Figure 1 schematically illustrates an example of an aircraft. Figure 2 is a schematic, partial, cross-sectional view of a twin-spool, turbofan aeronautical propulsion system in which the fan section is enclosed. Figure 3 is a schematic, partial, cross-sectional view of a twin-spool, turbofan aeronautical propulsion system in which the fan section is unenclosed.

[0031] Figure 4 schematically illustrates a blade of a stator part of a stage of a compressor of an aeronautical propulsion system, a device for controlling the angle of the blade pitch, and a data processing device for controlling the control device.

[0032] Figure 5 is a flowchart illustrating an example of the implementation of a process according to the present presentation.

[0033] Figure 6 is a flowchart illustrating an example of determining a target angle value for implementing a process according to this presentation.

[0034] Figure 7 is a flowchart illustrating another example of determining a target angle value for implementing a process as described herein. DETAILED DESCRIPTION

[0035] Aircraft

[0036] An aircraft is a device configured to rise and move through the air, and can, for example, be a civil or military airplane, or even a helicopter. An aircraft comprises an airframe (or "airframe" in Anglo-Saxon terminology) which, in the case of an airplane, consists of a fuselage, wings, tail assembly, control surfaces, and landing gear.

[0037] Propulsion system

[0038] A propulsion system 1 has a principal direction along a longitudinal axis X along which the propulsion system 1 extends. The propulsion system 1 is an aeronautical propulsion system 1 configured to be attached to the airframe of the aircraft 100 by means of a pylon (or mast), which is fixed to the airframe of the aircraft 100.

[0039] In this exposition, an axial direction corresponds to the direction of the longitudinal axis X, and a radial direction is a direction perpendicular to and passing through the longitudinal axis X. Furthermore, a circumferential (or lateral, or tangential) direction corresponds to a direction perpendicular to and not passing through the longitudinal axis X. Unless otherwise specified, the terms "internal" (or "inside") and "external" (or "outside") are used with reference to a radial direction such that the internal part or face of an element is closer to the longitudinal axis X than the external part or face of the same element.

[0040] The propulsion system 1 includes, from upstream to downstream in the direction of the gas flow in the propulsion system 1 when in operation, a blower section 2 and a generator, often called a "gas generator", which has a primary body 3. The primary body 3 is centered on the longitudinal axis X, and includes a compressor section 4, 5, a combustion chamber 6, and a turbine section 7, 8.

[0041] The blower section 2 includes at least one rotor 9 suitable for being driven in rotation, about the longitudinal axis X, relative to a stator portion of the propulsion system 1, by the generator and, more particularly, by at least one rotor portion 8a, 11 of the turbine section 7, 8 of the primary body 3. In this way, an airflow F is drawn into the propulsion system 1. Each rotor 9 of the blower section 2 includes a hub 13 and blades 14 extending radially from the hub 13. The blades 14 of each rotor 9 may be fixed relative to the hub 13 or have variable pitch. In this case, the base of the blades 14 of each rotor 9 is mounted pivoting about a pitch axis extending radially with respect to the longitudinal axis X. The base of the blades 14 is then connected to a piloting device 15 (or pitch changing mechanism 15) mounted in the propulsion system 1, the pitch being adjusted by the piloting device 15.

[0042] The blower section 2 may further include a stator 16, or rectifier, which comprises blades 17 mounted on a stator hub 16 and whose function is to rectify an airflow F2 exiting the rotor 9. The stator blades 17 may be fixed relative to the stator hub 16 or have variable pitch. If so, and similarly to the rotor blades 14 9, the base of the stator blades 16 is pivotally mounted about a radial pitch axis and is connected to a pilot device 15a (or pitch-changing mechanism 15a), which is generally separate from that of the rotor 9, the pitch being adjusted by the pilot device 15a. The stator hub 16 may be fixed to a stator portion of the generator.

[0043] The blower section 2 can be shrouded or unshrouded.

[0044] In the case of a shrouded blower section 2, the blower section 2 includes a blower housing 12, and the rotor 9 is housed in the blower housing 12. A shrouded blower section 2 includes a rotor 9 extending upstream of a stator 16. The blades 17 of the stator 16 are then generally called "outlet blades" (or "OGV", for "Outlet Guide Vane" in Anglo-Saxon terminology) and have a fixed position relative to the hub of the stator 16.

[0045] In the case of an unducted fan section 2, the fan section 2, which may also be referred to as the "propeller," is not enclosed by a fan casing. The blades 14 of the rotor 9 also have variable pitch. Propulsion systems 1 comprising at least one unducted rotor 9 are known, in Anglo-Saxon terminology, as "open rotor" or "unducted fan." The propulsion system 1 may comprise two unducted, counter-rotating rotors 9. Such a propulsion system 1 is known, in Anglo-Saxon terminology, by the acronym "CROR" for "Contra-Rotating Open Rotor" or "UDF" for "Unducted Double Fan." The rotors 9 can be placed at the rear of the generator so as to be of the pusher type or at the front of the generator so as to be of the tractor type.Alternatively, the propulsion system 1 may comprise a single unducted rotor 9 and an unducted stator 16 (rectifier). Such a propulsion system 1 is known, in Anglo-Saxon terminology, by the acronym "USF" for "Unducted Single Fan". In the case of a USF-type propulsion system 1, the blades 17 of the rectifier 16 are fixed in rotation relative to the stator portion of the propulsion system 1 and, consequently, are not subjected to centrifugal force. The blades 17 of the rectifier 16 are, moreover, of variable pitch. Removing the fairing around the fan section 2 allows a very significant increase in the dilution rate of the propulsion system 1 without the propulsion system 1 being penalized by the mass of the casings 12 or nacelles intended to surround the fan section 2. The compressor section 4, 5 comprises a succession of stages each comprising a rotating blade wheel (rotor) 4a, 5a rotating in front of a fixed blade wheel (stator) 4b, 5b.Each of the blades 40b, 50b of each stage of the fixed blade wheel 4b, 5b, in particular the blades 40b, 50b of the upstream stage, is pivotally mounted about a pitch axis Y extending radially with respect to the longitudinal axis X. In one variant, the upstream stage of the fixed blade wheel 4b of the compressor section 4, 5 is called the IGV (for "Inlet Guiding Vane" in Anglo-Saxon terminology), and is positioned upstream of the upstream stage of the rotating blade wheel 4a of the compressor section 4, 5. Furthermore, each of the blades 40b, 50b is controlled by a pilot device 400 mounted in the primary casing 3, and configured to adjust a pitch angle α of at least one of the blades 40b, 50b, if not each of the 40b, 50b blades, around its Y alignment axis. Preferably, the alignment angle thus adjusted is identical for all 40b, 50b blades.The turbine section 7, 8 also includes a succession of stages each comprising a fixed blade wheel (stator) 7b, 8b behind which rotates a moving blade wheel (rotor) 7a, 8a.

[0046] In a twin-spool propulsion system 1, the compressor section 4, 5 comprises a low-pressure compressor 4 and a high-pressure compressor 5, and the turbine section 7, 8 comprises a high-pressure turbine 7 and a low-pressure turbine 8. The rotor stages 5a of the high-pressure compressor 5 are driven in rotation by the rotor stages 7a of the high-pressure turbine 7 via a high-pressure shaft 10. The rotor stages 4a of the low-pressure compressor 4 and the rotor 9 of the blower section 2 are driven in rotation by the rotor stages 8a of the low-pressure turbine 8 via a low-pressure shaft 11. Thus, the primary body 3 comprises a high-pressure body including the high-pressure compressor 5, the high-pressure turbine 7, and the high-pressure shaft 10, and a low-pressure body including the blower section 2, the low-pressure compressor 4, the low-pressure turbine 8, and the low-pressure shaft 11.The rotational speed of the high-pressure body is greater than the rotational speed of the low-pressure body.

[0047] In a three-spool propulsion system 1, the turbine section 7, 8 further includes an intermediate turbine, positioned between the high-pressure turbine 7 and the low-pressure turbine 8, whose rotor stages are configured to drive the rotor stages of the low-pressure compressor 4 via an intermediate shaft. The rotor 9 of the blower section 2, on the one hand, and the rotor stages of the high-pressure compressor 5, on the other hand, remain driven by the low-pressure shaft 11 and the high-pressure shaft 10, respectively.

[0048] The low-pressure shaft 11 is generally housed, along a portion of its length, within the high-pressure shaft 10 and is coaxial with the high-pressure shaft 10. The low-pressure shaft 11 and the high-pressure shaft 10 may be co-rotating, that is, driven in the same direction around the longitudinal axis X. Alternatively, the low-pressure shaft 11 and the high-pressure shaft 10 may be counter-rotating, that is, driven in opposite directions around the longitudinal axis X. In any case, the high-pressure shaft 10 and the low-pressure shaft 11 each extend along the longitudinal axis X. If applicable, the intermediate shaft is housed between the high-pressure shaft 10 and the low-pressure shaft 11. The intermediate shaft and the low-pressure shaft 11 may be co-rotating or counter-rotating.

[0049] The generator includes an intermediate housing 33 (or inter-compressor housing 33) which supports the high-pressure shaft 10, the low-pressure shaft 11, and / or the intermediate shaft in rotation, via bearings. The intermediate housing 33 is fixed relative to the stator portion of the propulsion system 1.

[0050] The turbine section 7, 8 includes an intermediate casing 30 (or inter-turbine casing 30), positioned between the high-pressure turbine 7 and the low-pressure turbine 8, and a rear casing 31 (or outlet casing 31) positioned downstream of the low-pressure turbine 8. The intermediate casing 30 and the rear casing 31 are fixed relative to the stator part of the propulsion system 1. In addition, the intermediate casing 30 can support the rotation of the high-pressure shaft 10, while the rear casing 31 can support the rotation of the low-pressure shaft 11, via bearings.

[0051] The rotor 9 of the fan section 2 can be decoupled from the low-pressure shaft 11 by means of a reduction mechanism 19, positioned between an upstream end of the low-pressure shaft 11 and the rotor 9, in order to independently optimize their respective rotational speeds. In this case, the propulsion system 1 further includes an additional shaft (or fan shaft 20). The low-pressure shaft 11 connects the rotor stages 8a of the low-pressure turbine 8 to an inlet of the reduction mechanism 19, while the fan shaft 20 connects the outlet of the reduction mechanism 19 to the rotor 9 of the fan section 2. The rotor 9 of the fan section 2 is therefore driven by the low-pressure shaft 11 via the reduction mechanism 19 and the fan shaft 20 at a rotational speed lower than the rotational speed of the low-pressure turbine 8.This decoupling makes it possible to reduce the rotational speed and pressure ratio of the rotor 9 of the blower section 2 and to increase the power extracted by the low pressure turbine 8.

[0052] The reduction mechanism 19 may include an epicycloidal (or "planetary" in Anglo-Saxon terminology) or planetary (or "star" in Anglo-Saxon terminology) reduction mechanism, single-stage or two-stage.

[0053] The rotor 9 of the blower section 1 can, alternatively, be directly coupled to the low-pressure shaft 11 (or "direct-drive" in Anglo-Saxon terminology), i.e. without reduction mechanism 19. The low-pressure shaft 11 is then merged with the blower shaft 20 so that the rotor 9 is driven by the low-pressure shaft 11 at the same rotational speed as that of the rotor stages 8a of the low-pressure turbine 8.

[0054] The generator includes a front casing 32 (or inlet casing 32), which is positioned between the rotor 9 of the blower section 2 and the low-pressure compressor 4 when the propulsion system 1 is of the tractor type. The front casing 32 is fixed relative to a stator part of the propulsion system 1. The front casing 32 can support the rotation of the low-pressure shaft 11 and / or the blower shaft 20, via bearings. The reduction mechanism 19 can also be supported by the front casing 32.

[0055] The aeronautical propulsion system 1 further includes a data processing device 1000 (or "FADEC," for "Full Authority Digital Engine Control" in Anglo-Saxon terminology) comprising a processor. The data processing device 1000 is configured to control a number of elements of the aeronautical propulsion system 1, notably the piloting device 400 of the blades 40b, 50b of the fixed blade wheels 4b, 5b of the compressor section 4, 5.

[0056] In operation, an airflow F entering the propulsion system 1 is divided, by an annular nozzle 300 of the primary body 3, between a primary airflow F1 and a secondary airflow F2, which circulate from upstream to downstream in the propulsion system 1.

[0057] The primary airflow F1 flows in a primary channel 29 defined by the primary body 3, passing successively through the compressor section 4, 5, the combustion chamber 6 where it is mixed with fuel to serve as an oxidizer, and the turbine section 7, 8. The passage of the primary airflow F1 through the turbine section 7, 8 receiving energy from the combustion chamber 6 causes a rotation of the rotor stages 7a, 8a of the turbine section 7, 8, which in turn drives the rotation of the rotor stages 4a, 5a of the compressor section 4, 5 as well as a rotor 9 of the blower section 2.

[0058] The secondary airflow F2, also called the "bypass airflow", flows around the primary body 3. The secondary airflow F2 serves as a cold source for the propulsion system 1 and allows, in particular, the cooling of the periphery of the primary body 3. In addition, the secondary airflow F2 serves to generate most of the thrust provided by the propulsion system 1.

[0059] Steering the angle of adjustment

[0060] In a coordinate system providing a compression ratio of the low-pressure compressor 4 as a function of the primary flow rate F1 at the inlet of the low-pressure compressor 4, at a given altitude A of the aircraft propulsion system 1, and a given Mach number M of the aircraft 100, it is possible to plot field lines for each regime R of the low-pressure shaft 11, and field lines for each isentropic efficiency of the low-pressure compressor 4. The compression ratio of the low-pressure compressor 4 is defined as a ratio between a pressure within the primary flow rate F1 at the outlet of the low-pressure compressor 4, and a pressure within the primary flow rate F1 at the inlet of the low-pressure compressor 4. The primary flow rate F1 at the inlet of the low-pressure compressor 4 is defined as a ratio between the quantity of air in the primary flow rate F1 passing through the inlet of the low-pressure compressor 4 during a given time, and the given time.The iso-regime field lines are decreasing functions in the operating domain of the aeronautical propulsion system 1, while the isorythm field lines form concentric ellipses.

[0061] An operating line of the aeronautical propulsion system 1 can also be drawn in such a coordinate system. It connects each compression ratio / inlet flow rate pair for each operating regime R of the low-pressure shaft 11 of the aeronautical propulsion system 1 when it is in operation.

[0062] A pumping line of the low-pressure compressor 4 can also be drawn in such a coordinate system. Beyond the pumping line, the compression ratio of the low-pressure compressor 4 is sufficiently high, and the flow rate of the primary stream F1 at the inlet of the low-pressure compressor 4 is sufficiently low, that the low-pressure compressor 4 is subject to pumping, which is likely to lead to fatal damage to the low-pressure compressor 4, or even to the aircraft propulsion system 1. The aircraft propulsion system 1 is designed so that the concentric iso-efficiency lines are centered on a point positioned outside the pumping zone of the low-pressure compressor 4.The operating line of the aeronautical propulsion system 1 moves towards the pumping line as the aeronautical propulsion system 1 deteriorates, for example due to oxidation, erosion, and / or wear under extreme maneuvers of certain parts.

[0063] The aeronautical propulsion system 1 is designed so that the isentropic efficiency of the low-pressure compressor 4, and therefore the thermal efficiency of the aeronautical propulsion system 1, is maximized for a predetermined R speed of the low-pressure shaft 11, at a given altitude A of the aeronautical propulsion system 1, and a given Mach number M of the aircraft 100. Outside the optimal operating regime of the low-pressure compressor 4, the pitch angle a of the blades 40b of the fixed blade wheel 4b of the low-pressure compressor 4 can be adjusted by the piloting device 400, on command from the data processing device 1000, to maximize the isentropic efficiency of the low-pressure compressor 4 without however reaching pumping.In this case, at a given altitude A of the aeronautical propulsion system 1 and a given Mach number M of the aircraft 100, the data processing device 1000 commands the control device 400 to adjust the pitch angle α of the blades 40b of the fixed impeller 4b of the low-pressure compressor 4 according to a control law that depends on the speed R of the low-pressure shaft 11. At low speed R of the low-pressure shaft 11, the pitch angle α is controlled to be increased and thus limit the flow rate of the primary flow F1 at the inlet of the low-pressure compressor 4. This process is called "closing" the pitch. At high speed R of the low-pressure shaft 11, the pitch angle α is controlled to be decreased and thus increase the flow rate of the primary flow F1 at the inlet of the low-pressure compressor 4. This process is called "opening" the pitch.

[0064] At a given altitude A of the aeronautical propulsion system 1 and a given Mach number M of the aircraft 100, the control law for the blade pitch 40b of the fixed impeller 4b of the low-pressure compressor 4 generally does not vary according to the wear of the aeronautical propulsion system 1. Typically, this control law is determined so that, in the most deteriorated state of the aeronautical propulsion system, the surge zone is not reached. However, this penalizes the performance of the aeronautical propulsion system 1 since, in its undeteriorated state, it is thus kept away from areas of optimal isentropic efficiency of the low-pressure compressor 4, whereas its wear level would allow it to reach these areas without entering the surge zone.

[0065] Therefore, it is planned, within the framework of this presentation, to take into account the wear of the aeronautical propulsion system for the control of the pitch of the blades 40b of the fixed blade wheel 4b of the low pressure compressor 4. In this respect, the control law of the pitch of the blades 40b of the fixed blade wheel 4b of the low pressure compressor 4 is no longer fixed for the entire life of the aeronautical propulsion system 1, but varies according to its level of wear, while nevertheless retaining its overall shape of closing the pitch at low speed R of the low pressure shaft 11 and its opening of the pitch at high speed R of the low pressure shaft 11. In this way, areas of better isentropic efficiency of the low pressure compressor 4 can be reached when the engine is not completely degraded, without however reaching the pump line.

[0066] More specifically, a process E is implemented by the processor of the data processing device 1000 to take into account the wear of the aeronautical propulsion system 1. This process can be loaded into the data processing device 1000 in the form of a computer program product, which can also be stored on a readable storage medium.

[0067] This process E includes determining E1 a value for a target pitch angle ao of the blade 40b. This target pitch angle ao corresponds to the optimal pitch angle that the blade 40b must adopt to maximize the isentropic efficiency of the low-pressure compressor, at given altitude A and wear of the aeronautical propulsion system 1, given speed R of the low-pressure shaft 11, and given Mach number M of the aircraft 100, while avoiding surge of the low-pressure compressor 4. This value of the target pitch angle ao is therefore determined E1 from a value of a parameter P representative of a wear of the aeronautical propulsion system 1.

[0068] More specifically, the determination E1 of the target pitch angle value ao includes the determination of a control law L which provides values ​​of the target pitch angle Qo from values ​​of the low-pressure shaft R. This control law L usually depends only on the altitude A of the aircraft propulsion system 1 and the Mach number M of the aircraft 100. In the context of this exposition, the control law L also depends on the wear of the aircraft propulsion system 1, and more specifically on the value of the parameter P representing the wear of the aircraft propulsion system 1.Thus, for a given altitude A of the aeronautical propulsion system 1, and a given Mach number M of the aircraft 100, the process E makes it possible to determine, in a frame providing the target pitch angle ao as a function of the regime R of the low pressure shaft 1, a set of curves representing the successive control laws L, as a function of the wear of the aeronautical propulsion system 1. In this frame, the curve associated with the control law L for the pitch of the blade 40b in the new state of the aeronautical propulsion system 1 is positioned closer to the x-axis than the curve associated with the control law L for the aeronautical propulsion system 1 in its most degraded state.

[0069] Subsequently, the value of the target timing angle ao is determined E1 from a value of the regime R of the low pressure shaft 11, by taking the corresponding point of the control law L.

[0070] In one variant, the value of the parameter P representing the wear of the aeronautical propulsion system 1 is the number of operating cycles of the aeronautical propulsion system 1 since its commissioning. The number of cycles is readily available and thus allows for the simple determination E1 of the target pitch angle value ao. Furthermore, it is immediately apparent that the wear of the aeronautical propulsion system 1 increases with its use.

[0071] In another variant, the value of the parameter P representing the wear of the aeronautical propulsion system 1 is the difference between a value of the exhaust gas temperature T (or "EGT," for "Exhaust Gas Temperature" in Anglo-Saxon terminology) and a value of the exhaust gas temperature To at the start-up of the aeronautical propulsion system 1, at the same rotor speed Ro of the fan section 2. Indeed, as the aeronautical propulsion system 1 wears, the exhaust gas temperature T increases until it reaches a threshold of acceptability for the turbine section 7, 8, requiring the removal of the aeronautical propulsion system 1. The exhaust gas temperature T is the temperature of the primary flow F1 at the outlet of the combustion chamber 6. It is generally determined from the temperature of the primary flow F1 at the inlet of the low-pressure turbine 8.The value of the exhaust gas temperature To during the commissioning of the aeronautical propulsion system 1 can be determined from a reference law Lo providing values ​​of the exhaust gas temperature To during the commissioning of the aeronautical propulsion system 1 as a function of the speed Ro of the rotor 9 of the fan section 2. Indeed, it has been observed that the evolution of the exhaust gas temperature T is monotonic as a function of the speed Ro of the rotor 9 of the fan section 2, the corresponding curve simply shifting upwards in a reference frame providing the exhaust gas temperature T as a function of the speed Ro of the rotor 9 of the fan section 2.Therefore, the exhaust gas temperature To during the commissioning of the aeronautical propulsion system 1 is determined from the reference law Lo and a value of the rotor speed Ro of the fan section 2, by noting the corresponding point of the reference law Lo. Using the difference between the exhaust gas temperature T and the exhaust gas temperature To during the commissioning of the aeronautical propulsion system 1 as the value of the parameter P representing the wear of the aeronautical propulsion system 1 provides a more precise view of the wear of the aeronautical propulsion system 1.

[0072] Once the target pitch angle value ao is determined E1, the process includes generating E2 a control signal S configured to command the pilot device 400 so that it adjusts the pitch angle of the blade 40b to the determined target pitch angle value cto. The control signal S is then transmitted to the pilot device 400, which adjusts the pitch angle a of the blade 40b accordingly.

[0073] Although process E was described with reference to the timing of the fixed blades 40b of the low-pressure compressor 4, it is of course applicable to any system allowing variation of the flow rate of the primary flow F1 which, thanks to process E, would also take into account the wear of the propulsion system 1.

Claims

DEMANDS 1. A computer-implemented method (E) comprising: the determination (E1) of a value of a target pitch angle (ao) of a blade (40b, 50b) of a stator part of a stage of a compressor (4, 5) of an aeronautical propulsion system (1) of an aircraft (100) from a value of a parameter (P) representative of wear of the aeronautical propulsion system (1); and the generation (E2) of a control signal (S) configured to control a blade (400) pilot device (40b, 50b) so that the pilot device (400) adjusts a pitch angle (a) of the blade (40b, 50b) to the determined value of the target pitch angle (Cio).

2. Method (E) according to claim 1, wherein the determination (E1) of the value of the target pitch angle (ao) is further carried out from a value of an altitude (A) of the aeronautical propulsion system (1).

3. Method (E) according to any one of claims 1 and 2, wherein the determination (E1) of the value of the target pitch angle (ao) is further carried out from a value of a Mach number (M) of the aircraft (100).

4. Method (E) according to any one of claims 1 to 3, wherein the determination (E1) of the value of the target pitch angle (ao) is further carried out from a value of a regime (R) of the aeronautical propulsion system (1).

5. Method (E) according to any one of claims 1 to 4, wherein the determination (E1) of the value of the target pitch angle (ao) comprises the determination of a control law (L) providing values ​​of the target pitch angle (ao) from values ​​of a regime (R) of the aeronautical propulsion system (1); in which the control law is determined from the value of the parameter (P) representing the wear of the aeronautical propulsion system (1), a value of an altitude (A) of the aeronautical propulsion system (1), and a value of a Mach number (M) of the aircraft (100); in which the value of the target pitch angle (ao) is determined from the control law and a value of the regime (R) of the aeronautical propulsion system (1).

6. Method (E) according to any one of claims 1 to 5, wherein the value of the parameter (P) representing the wear of the aeronautical propulsion system (1) is a number (N) of operating cycles of the aeronautical propulsion system (1) since its commissioning.

7. Method (E) according to any one of claims 1 to 5, wherein the determination (E1) of the value of the target pitch angle (ao) comprises the determination of a temperature value (To) of the exhaust gases of the aeronautical propulsion system (1) when it is put into service; in which the value of the temperature (To) of the exhaust gases of the aeronautical propulsion system (1) when it is put into service is determined from a value of a regime (Ro) of the aeronautical propulsion system (1), and a reference law (Lo) providing values ​​of the temperature (To) of the exhaust gases of the aeronautical propulsion system (1) when it is put into service as a function of values ​​of the regime (Ro) of the aeronautical propulsion system (1); in which the value of the parameter (P) representing the wear of the aeronautical propulsion system (1) is a difference between a value of a temperature (T) of the exhaust gases of the aeronautical propulsion system (1) and the determined value of the temperature (To) of the exhaust gases of the aeronautical propulsion system (1) when it was put into service.

8. Product computer program comprising instructions which, when the program is executed by a computer, cause the computer to implement the method (E) according to any one of claims 1 to 7.

9. Data processing device (1000) for an aeronautical propulsion system (1) comprising a processor configured to implement the method (E) according to any one of claims 1 to 7.

10. Aeronautical propulsion system (1) having a longitudinal axis (X) along which the aeronautical propulsion system (1) extends, the aeronautical propulsion system (1) comprising: the data processing device (1000) according to claim 9; a compressor (4, 5) comprising a stage, the stage comprising a stator part (4b, 5b), the stator part (4b, 5b) comprising a blade (40b, 50b) having a pitch axis (Y) along which the blade (40b, 50b) extends, the pitch axis (Y) extending radially with respect to the longitudinal axis (X), the blade (40b, 50b) being configured to pivot about the pitch axis (Y) so as to have a pitch angle (a); and a control device (400) configured to adjust the pitch angle (a) of the blade (40b,50b) to the value of the target pitch angle (cto) from the control signal (S) generated by the data processing device (1000).

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