Method for maintaining the operation of an aircraft engine in the event of a pressure measurement failure

Abstract

One aspect of the invention relates to a method for compensating for a pressure measurement failure in an aircraft engine, the method including: - detecting (110) a pressure measurement failure by: o detecting a pressure measurement anomaly by comparing (111) the pressure measured at the outlet of a high-pressure compressor of the engine with a pressure model; and o detecting (112) an inconsistency between the operating speed of the aircraft engine and a setpoint; - compensating (120) for the measurement failure by: o replacing (121) the pressure measurement with a pressure obtained by the pressure model; o determining (122) a surge protection limit and a flameout protection limit on the basis of a time derivative of the operating speed, the surge and flameout protection limits replacing erroneous protection limits caused by the pressure measurement anomaly; o determining (123) a low-power flameout limit on the basis of an idling speed of a high-pressure section of the engine and / or a predefined limit of the pressure at the outlet of the high-pressure compressor, the low-power flameout limit being added to the surge and flameout protection limits.

Description

DESCRIPTION TITLE: METHOD FOR MAINTAINING THE OPERATION OF AN AIRCRAFT ENGINE IN THE EVENT OF A PRESSURE MEASUREMENT FAILURE TECHNICAL FIELD OF THE INVENTION

[0001] The technical field of the invention is that of controlling the speed of an aircraft.

[0002] The present invention relates to a method for compensating for a pressure measurement failure in an aircraft engine. TECHNOLOGICAL BACKGROUND OF THE INVENTION

[0003] In an aircraft, the pilot's flight instructions are transmitted to a computer, called FADEC (Full Authority Digital Engine Control), which is responsible for translating these instructions into commands for the aircraft's various operating components. In this computer, these instructions are, in particular, compared with navigation instructions and parameters specific to each aircraft. These navigation instructions and parameters, such as maximum and / or minimum fuel consumption, are, for example, evaluated based on data from measurements in the engine.

[0004] In particular, the engine's operating speed, i.e. its rotational speed, is limited by safety instructions relating to a maximum flow rate and a minimum flow rate of fuel that can be supplied to the engine. These maximum and minimum flow rates are calculated from measurements of physical properties, including the pressure downstream of the high-pressure compressor. The pressure at this location is conventionally obtained by calculating a static pressure, by the computer, in a thin tube, called a "capillary", one end of which is connected to the measurement location and the other to the computer.

[0005] The measurement of this pressure is however altered by the extreme environmental conditions of the device, for example in the event of the presence of water or frost in the tube, which distort the pressure calculated by the calculator. Devices to protect the tubes from these extreme conditions, for example to defrost, exist but remain insufficient and may not detect a measurement anomaly for this pressure, risking loss of control of the device.

[0006] There is therefore a need to improve the control of an aircraft in the event of an anomaly in the measurement of the pressure at the outlet of the high pressure compressor. SUMMARY OF THE INVENTION

[0007] The invention provides a solution to the problems mentioned above, by making it possible to detect a measurement failure not avoided or compensated for by the approaches of the art and which makes it possible to implement a mechanism for maintaining control of the device.

[0008] A first aspect of the invention relates to a method for compensating for a pressure measurement failure in an aircraft engine, the method comprising: Detect a pressure measurement failure by: Detection of a pressure measurement anomaly by comparing the pressure measured at the outlet of a high-pressure compressor of the engine with a pressure model; and Detection of an inconsistency in the operating speed of the aircraft engine with a setpoint; Compensate for measurement failure by: Substitution of the pressure measurement by a pressure obtained by the pressure model; Determination of a pumping protection limit and a shutdown protection limit from a time derivative of the operating regime, the pumping and shutdown protection limits replacing erroneous protection limits due to the pressure measurement anomaly; Determination of a low power extinguishing limit from an idle speed of a high pressure body of the engine and / or a predefined limit of the pressure at the outlet of the high pressure compressor, the low power extinguishing limit being added to the pumping and extinguishing protection limits.

[0009] "Compensation" means a mechanism or logic, for example implemented in the form of instructions in a computer, which makes it possible to replace erroneous information, in this case a pressure measurement, by replacing this information with one or more alternative data (the pressure modeled by the pressure model, the surge protection limit, the shutdown protection limit and the low power shutdown limit). These alternative data are then used to correct the piloting instructions for the aircraft components produced by the computer, particularly concerning the engine speed.

[0010] A "failure" in pressure measurement means an error in the calculation of pressure by the computer due to a disturbance in the measurement device, thus producing an erroneous pressure value compared to the actual pressure at the measurement point, i.e. the pressure at the outlet of the high-pressure (HP) compressor. In this case, the error in the calculation of pressure by the computer may be the result of an obstruction in the capillary carrying the pressure from the outlet of the HP compressor to the computer, for example due to the presence of water or ice in this capillary.

[0011] A "measurement anomaly" is a contradiction between the measured pressure and the pressure estimated by the pressure model. In particular, there is an anomaly when there is a significant distance between the measurement and the model.

[0012] An "inconsistency" in the operating speed is understood to mean a contradiction between the speed at which the engine is operating and the instruction sent to it by the computer. This inconsistency is linked to the measurement anomaly mentioned above. That is to say, the behavior of the engine is not consistent with the instruction determined by the computer to respond to the pilot's instruction because this instruction is based on an erroneous measurement of the pressure downstream of the HP compressor. An instruction is therefore one or more commands issued by the computer to indicate to the engine components how they must operate or modify their operating mode to satisfy the instruction provided by the pilot.

[0013] The terms "surge protection limit" and "flame-out protection limit" are two maximum and minimum limit values ​​for the fuel flow supplied to the engine to prevent the engine from surging or flaming. These values ​​are used to replace the limits conventionally calculated by the computer since the latter are erroneous due to the pressure measurement anomaly.

[0014] A "low power shutdown limit" is a limit value to prevent the engine from shutting down when it is in a low power operating mode. This value ensures that the engine will not have extinction even if the engine enters a low power operating mode and despite the failure of the pressure measurement.

[0015] Thanks to the invention, it is thus possible to detect a failure in the pressure measurement and then trigger logic to maintain control of the aircraft.

[0016] The measurement failure is detected by determining two indicators: a significant difference between the measured pressure and the pressure model, this model allowing the temperature at the HP compressor outlet to be theoretically evaluated; and an inconsistency between the engine's behavior and the instruction sent to it by the computer, which is determined based on the pressure measurement which is erroneous.

[0017] Compensation is then implemented to compensate for the pressure measurement failure. Compensation involves replacing the pressure measurement with the theoretical pressure calculated by the pressure model, which ensures that the computer constructs its engine control commands in accordance with the effective pressure of the HP compressor outlet, i.e. the actual pressure downstream of said compressor and not the abnormal pressure. Compensation also includes determining three limits to replace those conventionally calculated by the computer and used to define the engine operating limits. Indeed, the conventionally calculated limits are erroneous here since the pressure measurement does not conform to the actual pressure in the engine. Compensation therefore replaces these values ​​with values ​​that ensure that the engine operates without major malfunction, in particular without the risk of surge and / or flameout.

[0018] In other words, compensation allows the engine control to be reconfigured via operating logic to ensure that the pilot retains control of the aircraft even though the compressor outlet pressure measurement is faulty.

[0019] Thus, the pilot retains control of his aircraft even though there is a failure in the measurement of the pressure at the HP compressor outlet. Compensation can continue until the failure is resolved, for example by a capillary protection device or until the aircraft is stopped, once its flight is over.

[0020] Furthermore, the proposed method allows not to be sensitive to a common mode failure. Indeed, a redundancy of the pressure measurement can be implanted and also be faulty due to the fact that the pressure is conveyed by the same capillary which is duplicated at the level of the redundant computers.

[0021] Finally, and in the case of measurement redundancy, the proposed method makes it possible to compensate for the failure and maintain control of the aircraft even if a double sensor failure occurs, i.e. the respective pressure-sensitive sensors of the two redundant computers are both faulty due to an anomaly other than a fault in the capillary.

[0022] In addition to the characteristics which have just been mentioned, the method according to the first aspect of the invention may have one or more additional characteristics among the following, considered individually or according to all technically possible combinations.

[0023] In one embodiment, the method according to the invention comprises: Modifying the engine speed from the pressure obtained by the pressure model, the surge protection limit, the flameout protection limit and the low power flameout limit.

[0024] The engine operating speed is then corrected to comply with the instruction provided by the pilot, following new or corrected instructions produced by the computer.

[0025] In one embodiment, the measurement anomaly is detected when a deviation of a measured pressure value from a theoretical value determined by the pressure model is greater than a predefined threshold.

[0026] In one embodiment, a measured value is obtained by measuring the pressure and a theoretical value is determined by the pressure model, and the substitution of the pressure measurement is implemented by replacing the measured value with the theoretical value.

[0027] In one embodiment, the setpoint is a fuel flow rate and the inconsistency is detected when: the engine operating speed corresponds to deceleration or steady state while the fuel flow rate is equal to a maximum flow rate; or the engine operating speed corresponds to acceleration while the fuel flow rate is equal to a minimum flow rate.

[0028] In other words, engine operating speed inconsistency is detected when: Although the computer issues a command to reach the maximum fuel flow, i.e. the engine speed is expected to accelerate, the engine speed remains unchanged or decelerates; or Although the computer issues a command to reach the minimum fuel flow, i.e. it is expected that the engine speed will decelerate, the engine speed will accelerate.

[0029] A second aspect of the invention relates to a device configured to implement the method according to the first aspect.

[0030] A third aspect of the invention relates to an aircraft comprising the device according to the second aspect of the invention.

[0031] A fourth aspect of the invention relates to a computer program product comprising instructions which, when the program is executed on a computer, cause the latter to implement the steps of the method according to the first aspect.

[0032] A fifth aspect of the invention relates to a computer-readable recording medium comprising instructions which, when executed by a computer, cause the computer to carry out the steps of the method according to the first aspect.

[0033] The invention and its various applications will be better understood by reading the following description and examining the accompanying figures. BRIEF DESCRIPTION OF THE FIGURES

[0034] The figures are presented for information purposes only and in no way limit the invention. Figure 1 is a block diagram illustrating the sequence of steps of a method for compensating for a pressure measurement failure in an aircraft engine, according to one embodiment. Figure 2 is a schematic representation of a device configured to implement a method of compensating for a failure to measure pressure in an aircraft engine, according to one embodiment. Figure 3 is a diagram illustrating the determination of a pressure estimate by a theoretical model according to one embodiment. Figure 4 is a block diagram illustrating the implementation according to one embodiment of the method for compensating for a pressure measurement failure in an aircraft engine. DETAILED DESCRIPTION

[0035] Unless otherwise specified, the same element appearing in different figures has a single reference.

[0036] As described below, the invention proposes a method for compensating for a pressure measurement failure in an aircraft engine. This method makes it possible to substitute erroneous data processed by a computer with corrected or alternative data. The computer, also known as FADEC for "Full Authority Digital Engine Control" in English, is responsible for translating the pilot's piloting instructions into instructions for the components of an aircraft engine. The method comprises two main steps: the first is used to detect the pressure measurement failure and the second to compensate for the failure.

[0037] The sequence of steps of the method 100 is illustrated in Figure 1.

[0038] Step 110 of detecting a pressure measurement failure is first implemented. This detection is based on the evaluation of two indicators representative of this failure, namely a comparison of the measured pressure with a theoretical model, and the identification of inconsistency between the actual operation of the engine and the instruction emitted by the computer.

[0039] Step 110 therefore includes a step 111 of comparing the measured pressure with the pressure model.

[0040] The pressure is measured at the outlet of the high-pressure (HP) compressor of the aircraft engine. The measurement is carried out using conventional techniques, for example via a capillary connecting the measurement location to the computer. The measured pressure is a static or quasi-static pressure.

[0041] The measurement is subject to malfunction here, for example due to the presence of frost or water in the capillary or due to a malfunction of the pressure-sensitive element of the computer connected to the capillary.

[0042] The pressure model is a predefined theoretical model, for example a known model or a model established by an operator. The model can therefore be any model that allows estimating a theoretical pressure that the computer should detect via its sensitive element. Advantageously, the pressure model can be a simplified or approximate model, with a margin of error less than or equal to 30%.

[0043] The measurement anomaly is therefore detected when a difference between the measured pressure and the pressure estimated by the model is greater than a predefined threshold. The predefined threshold is, for example, established by an operator based on the type and model of the aircraft, the characteristics of the computer and / or the engine components, and / or their professional knowledge.

[0044] In particular, a pressure value, called the "measured value", is obtained from the pressure measurement. Similarly, an estimated pressure value, called the "theoretical value", is obtained from the pressure model. Thus, the measurement anomaly is detected when a deviation between the measured value and the theoretical value is greater than the predefined threshold. The deviation between these two values ​​is, for example, equal to the absolute value of a difference between these two values. Other types of deviation can be used, such as an absolute value of a difference of the squares of the values, a summation of the values, a summation of the square of the values, etc.

[0045] Step 110 also includes a step 112 for detecting the inconsistency between the actual operation of the engine and the setpoint issued by the computer. In particular, this step is used to detect an inconsistency between the operating speed of the engine and the setpoint. The operating speed of the engine is here the rotation speed of the shafts N1 and / or N2.

[0046] An inconsistency can be detected when the engine speed decelerates or remains at a steady speed while the setpoint indicates that the engine speed should accelerate. Alternatively, an inconsistency can be detected when the engine speed accelerates while the setpoint indicates that the engine speed should decelerate.

[0047] In particular, the instruction issued by the computer relates to a fuel flow rate that must be supplied to the engine to satisfy the instruction given by the pilot. The fuel flow rate indicated by the instruction and determined by the computer is limited by a maximum threshold and a minimum threshold, preventing the engine from entering surge or flameout. Consequently, when the fuel flow rate indicated in the instruction is equal to the maximum threshold or the minimum threshold, the engine speed is expected to accelerate or decelerate, respectively.

[0048] Therefore, the inconsistency between the engine operating speed and the setpoint is detected when: The engine operating speed is at deceleration or steady state while the engine fuel flow rate is at the maximum threshold; or The engine operating speed corresponds to acceleration while the fuel flow is equal to the minimum threshold.

[0049] The maximum and minimum thresholds are determined by the calculator, in a manner known per se.

[0050] Therefore, when the measurement anomaly and the operating regime inconsistency are detected, then the measurement failure is detected.

[0051] Step 120 of failure compensation is then implemented. This step comprises three steps described below and relating to the substitution of the data necessary for the computer to determine the instructions by alternative or replacement data.

[0052] Step 120 thus includes a step 121 of substituting the measured pressure with the pressure estimated by the pressure model. In other words, the measured pressure value is replaced, at the computer level, by the theoretical value. The computer can thus use a pressure value representative of the actual pressure at the outlet of the HP compressor to construct the setpoints, in particular to indicate the fuel flow to be supplied to the engine to satisfy the pilot's instruction.

[0053] Step 120 also includes a step 122 for determining a pumping protection limit and a flameout protection limit. These limits are used to condition the computer so that the fuel flow rate that it calculates and indicates in its setpoint does not exceed the pumping protection limit and is not less than the flameout protection limit. When there is no failure, these limits are calculated by the computer using techniques known per se. However, in the event of a failure, the conventional approaches fail and cause the inconsistency mentioned above. Consequently, the two determined limits are substituted for the conventionally determined limits, at the computer level. The computer therefore takes these determined limits into account and no longer the conventionally determined limits, to construct its setpoint and calculate the related fuel flow rate.

[0054] The pumping protection and flameout protection limits are determined from the operating regime, in particular from a time derivative of the operating regime. For example, the time derivative is that of the rotational speed of shaft N1 or shaft N2.

[0055] These limits are more restrictive than the conventionally calculated limits. They therefore ensure that the engine does not surge or flame out and that the pilot retains control of the aircraft.

[0056] Step 120 finally includes a step 123 for determining a low power shutdown limit. This limit indicates a value at which the fuel flow rate must not be lower than the risk of causing the engine to shut down while it is in low power mode. This limit is therefore added to the limits calculated in step 122 to constrain the computer in its determination of the setpoint and the calculation of the related fuel flow rate.

[0057] The low power extinction limit is determined from a high pressure body speed of the engine, i.e. the rotational speed of said body. The low power extinction limit can, alternatively or jointly, be calculated from a predefined limit of the pressure at the outlet of the high pressure compressor. This predefined limit is defined by the pressure model and / or by the operator depending on the type and model of the aircraft, the characteristics of the computer and / or the engine components, and / or his professional knowledge.

[0058] The method 100 may also comprise a step 130 of modifying the operating speed of the engine. During this step, the computer determines a new setpoint, or a corrected setpoint, and therefore a new fuel flow rate taking into account the replacement data determined in step 120.

[0059] The engine speed is therefore modified from the pressure obtained by the pressure model, the pumping protection limit, the flameout protection limit and the low power flameout limit.

[0060] The method 100 can be implemented by a device for compensating for a failure in measuring pressure in an aircraft engine. Said device is therefore configured to implement said method 100. Such a device 10 comprises, for example and as shown diagrammatically in FIG. 2, a circuit comprising a processor 11, a volatile or non-volatile memory 12. The memory 12 is capable of storing instructions which, when implemented by the processor 11, lead the processor 11 to implement the steps of the method 100. The device 10 is, for example, a computer.

[0061] The circuit may alternatively comprise an electronic card whose steps of the method of the invention are described in the silicon, or even a programmable electronic chip such as an FPGA chip (for “Field-Programmable Gate Array” in English).

[0062] The device 10 also comprises an input interface 13 and an output interface 14. The input interface 13 is, for example, capable of receiving the pressure measurement to be compared to the model. More particularly, the input interface 13 is capable of receiving the measured value of the pressure. The input interface 12 is also capable of receiving the instruction emitted by the computer and the operating speed in which the engine is located. In particular, the input interface 12 is capable of receiving data indicating the operating speed of said engine.

[0063] The output interface 14 is capable of the theoretical pressure value obtained by the pressure model, the pumping protection limit, the extinguishing protection limit and the low power extinguishing limit determined via the implementation of the method 100.

[0064] The device 10 may, furthermore, be configured to implement calculations of the pressure model used by the method 100 and mentioned above. The device 10 therefore comprises, in its memory 12, instructions which, when implemented implemented by the processor 11, lead the latter to implement calculations of the pressure model. Alternatively, the calculations of the pressure model are implemented by an additional module dedicated to these calculations. This module is thus configured to provide the device 10 with the pressure estimated by the model, i.e. the theoretical value, in which case the input interface 13 of the device 10 is also capable of receiving this estimated pressure.

[0065] In some embodiments, the device 10 is attached to the computer. In other embodiments, the device 10 is included in the computer.

[0066] When two computers are used in redundancy, the device 10 may be a single device 10 for both computers or may also be in redundancy, i.e. there is a device 10 for or in each computer.

[0067] For the purpose of illustration, an example of implementation of the method 100, in particular step 110, is detailed below with reference to figures 3 and 4.

[0068] The pressure measurement is carried out using a capillary connected on one side to the HP compressor outlet and on the other side to a pressure sensor in the aircraft's computer or FADEC.

[0069] The method 100 is implemented on a device 10 such as that presented above.

[0070] The device 10 obtains the measured pressure value from the computer and compares it to the theoretical value estimated via the following pressure model. This model is an example of simplified modeling of the pressure in a turbomachine and has the advantage of being simple and quick to implement. This model is suitable for twin-spool dual-flow turbojets.

[0071] As illustrated in Figure 3, obtaining the theoretical pressure value is determined by multiplying data from two different sub-models: The first sub-model M1 is configured to determine a first pressure as a function of the engine's low pressure compressor speed; and The second sub-model M2 is configured to determine a second pressure based on the high pressure compressor speed.

[0072] The first sub-model M1 takes as input a value of the low pressure compressor speed, i.e. its rotation speed, expressed in percentage of the maximum speed of the high pressure compressor, and noted PCN12R. In the model, the speed of the low pressure compressor, noted X n 12 is used to determine X n 12R such that X n 12R = where T 12is the estimated or measured temperature 7'12 at the outlet of the low pressure compressor. The value of PCN12R is expressed as a function of X n 12R, for example in the form of a ratio with the maximum speed of the high pressure compressor noted X n 25MAX.

[0073] The second model M2 takes as input a value of the high pressure compressor speed, i.e. its rotation speed, expressed as a percentage of the maximum speed of this high pressure compressor, and noted PCN25R. In the model, the speed of the low pressure compressor, noted X n 25 is used to determine X n 25R such that X n 25R = where T 25 is the estimated or measured temperature at the outlet of the low pressure compressor. The value of PCN25R is expressed as a function of X n 25R, for example as a ratio with X n 25MAX.

[0074] The sub-models M1 and M2 are known per se and are, for example, obtained by analyses and measurements carried out during engine tests according to approaches known to those skilled in the art.

[0075] The first pressure calculated by the first model M1 is multiplied by an estimate, or measurement, of the total pressure PT2 upstream of the rotor of the low-pressure compressor. This multiplication produces an intermediate pressure value. This intermediate value is then multiplied by the second calculated pressure. At the end of the second multiplication, the theoretical value PS3 is produced. th of the pressure estimated by the pressure model.

[0076] The theoretical value of PS3 th is then compared, in step 111 and as illustrated in Figure 4, to the measured value, which is noted PS3. It is considered that the precision of this model is equal to 30% and is noted e 30The measurement anomaly is thus determined when: PS3 < PS3 th (l - e); or when PS3 > PS3 th (1 + e).

[0077] The error e is the approximation error of the pressure model. For the model presented, the error here is 30%.

[0078] In the proposed example, at step 112 and as illustrated in Figure 4, the inconsistency is detected when: WF32 cmd = WF32 max and that when - WF32 cmd = WF32 min and that

[0079] The fuel flow indicated in the calculator instruction is noted WF32 cmd The maximum fuel flow allowed for the engine, also determined by the computer, is noted WF32 rnax The minimum fuel flow allowed for the engine, also determined by the computer, is noted WF32 min. The time derivative of the rotational speed of shaft N2 is noted where t is time.

[0080] It is noted that the inconsistency could be indifferently detected by using the time derivative of the speed of shaft N1 instead of that of shaft N2.

[0081] It is possible to perform the inconsistency detection redundantly, for example a second time at a time following the first inconsistency detection to confirm this detection.

[0082] The measurement failure is then detected in two cases: Where the inconsistency is due to the N2 shaft speed indicating deceleration or stagnation while the control fuel flow rate is equal to the maximum fuel flow rate and the measured pressure value is less than the theoretical pressure value less the error; or When the inconsistency is due to the fact that the N2 shaft speed indicates acceleration while the control fuel flow is equal to the minimum fuel flow and the measured pressure value is greater than the theoretical pressure value plus the error.

[0083] In addition to the conditions of these two cases, the failure can be detected provided that the engine is not stalled, that is, the engine is not pumping, and provided that said engine is switched on.

[0084] Step 120 is then implemented by the device and, in step 121, the theoretical pressure value is used by the computer to determine the setpoints instead of the measured value.

[0085] In step 122, the pumping protection limit and the flameout protection limit are calculated and replace the maximum and minimum fuel flow rates conventionally calculated by the computer such as: WF32 max = (1 + e). C / P max . PS3 th. / T25 / b ; and - WF32 min = (1 - e C / P max .PS3 th . r23 / b.

[0086] The value of the coefficient b is determined by the operator and is, for example, b = 288.15. C!P max is a high limit that limits the fuel flow to protect the HP compressor from surging due to overconsumption of surging margin related to thermal throttling created by the fuel level. The value of C / P-max can be determined by any known state-of-the-art method.

[0087] In addition, control loops are implemented and serve to limit the acceleration rate and the deceleration rate of the HP body. The acceleration rate is thus limited by a maximum value noted and a minimum value noted These regulation loops, known in themselves, allow to apply more restrictive protection but nevertheless allowing acceleration and deceleration to be ensured in the event of pressure measurement failure.

[0088] As a result, the engine is doubly protected from surge or flameout by overriding the maximum and minimum fuel flow rates and by limiting the acceleration and deceleration rates.

[0089] The device can then transmit this data to the computer which can construct, according to step 130, then send to the devices of the device a new instruction based on this data to respond to the piloting instruction required by the pilot.

Claims

CLAIMS

1. A method (100) of compensating for a failure to measure pressure in an aircraft engine, the method comprising: - Detecting (110) a pressure measurement failure by: o Detecting a pressure measurement anomaly by comparing (111) the pressure measured at the outlet of a high-pressure compressor of the engine with a pressure model; and o Detecting (112) an inconsistency in the operating speed of the aircraft engine with a setpoint; - Compensating (120) the measurement failure by: o Substituting (121) the pressure measurement with a pressure obtained by the pressure model; o Determining (122) a surge protection limit and a shutdown protection limit from a time derivative of the operating regime, the surge and shutdown protection limits replacing erroneous protection limits due to the pressure measurement anomaly; o Determining (123) a low power shutdown limit from an idle speed of a high pressure body of the engine and / or a predefined limit of the pressure at the outlet of the high pressure compressor, the low power shutdown limit being added to the surge and shutdown protection limits.

2. Method according to the preceding claim, comprising: - Modify (130) the engine speed from the pressure obtained by the pressure model, the pumping protection limit, the flameout protection limit and the low power flameout limit.

3. Method according to one of the preceding claims, wherein the measurement anomaly is detected when a deviation of a measured value of the pressure with a theoretical value determined by the pressure model is greater than a predefined threshold.

4. Method according to one of the preceding claims, a measured value being obtained by measuring the pressure and a theoretical value being determined by the pressure model, wherein the substitution of the pressure measurement is implemented by replacing the measured value with the theoretical value.

5. Method according to one of the preceding claims, in which the setpoint is a fuel flow rate and in which the inconsistency is detected when: - The engine operating speed corresponds to deceleration or steady state while the fuel flow rate is equal to a maximum threshold; or - The engine operating speed corresponds to acceleration while the fuel flow is equal to a minimum threshold.

6. Device (10) for compensating for a failure to measure a pressure in an aircraft engine configured to implement the method (100) according to one of the preceding claims.

7. Aircraft comprising the device (10) according to the preceding claim.

8. A computer program product comprising instructions which, when the program is executed on a computer, cause the latter to implement the steps of the method (100) according to one of claims 1 to 5.

9. A computer-readable recording medium comprising instructions which, when executed by a computer, cause the computer to carry out the steps of the method (100) according to one of claims 1 to 5.

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