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

Abstract

Method for maintaining the operation of an aircraft engine in the event of a pressure measurement failure. One aspect of the invention relates to a method for compensating for a pressure measurement failure in an aircraft engine. Figure to be published with the abbreviation: Figure 4

Description

Title of the invention: 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 aircraft regime control.

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

[0003] In an aircraft, the flight instructions given by the pilot during flight are transmitted to a computer, called FADEC for "Full Authority Digital Engine Control," which translates these instructions into commands for the various operating systems of the aircraft. 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 engine measurements.

[0004] In particular, the engine's operating regime, i.e., its rotational speed, is limited by safety instructions concerning a maximum and a minimum fuel flow rate 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, using the computer, in a thin tube, called a "capillary tube," one end of which is connected to the measurement point and the other to the computer.

[0005] However, the measurement of this pressure is affected by extreme environmental conditions in the device, for example, the presence of water or frost in the tube, which distorts the pressure calculated by the computer. Devices to protect the tubes from these extreme conditions, for example, defrosting devices, exist but remain insufficient and may fail to detect a measurement anomaly for this pressure, potentially causing the device to malfunction.

[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 offers 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 to maintain 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: • Detecting a pressure measurement failure by: • Detection of a pressure measurement anomaly by comparing the pressure measured at the outlet of a high-pressure engine compressor with a pressure model; and • Detection of an inconsistency between the aircraft engine's operating mode and a setpoint; • Compensate for the measurement failure by: • Substitution of the pressure measurement with a pressure obtained from the pressure model; • Determination of a pumping protection limit and an extinction protection limit from a time derivative of the operating regime, the pumping and extinction protection limits replacing erroneous protection limits due to the pressure measurement anomaly; • Determination of a low power extinction 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 extinction limit being added to the pumping and extinction protection limits.

[0009] The term "compensation" refers to a mechanism or logic, for example implemented as instructions in a computer, that allows erroneous information, in this case a pressure measurement, to be replaced by one or more alternative data points (the pressure modeled by the pressure model, the pumping protection limit, the extinction protection limit, and the low-power extinction limit). These alternative data points are then used to correct the aircraft control commands generated by the computer, particularly regarding engine speed.

[0010] A "pressure measurement failure" is defined as an error in the pressure calculation by the computer due to a disturbance in the measuring device, resulting in 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 pressure calculation by the computer may be the result of an obstruction in the capillary tube carrying the pressure from the HP compressor outlet to the computer, for example, due to the presence of water or frost in this capillary tube.

[0011] A "measurement anomaly" is understood to mean a contradiction between the measured pressure and the pressure estimated by the pressure model. In particular, an anomaly exists when the difference between the measurement and the model is significant.

[0012] The term "inconsistency" in the operating mode refers to a contradiction between the engine's operating speed and the command sent to it by the engine control unit (ECU). This inconsistency is linked to the measurement anomaly mentioned above. In other words, the engine's behavior is not consistent with the command determined by the ECU to respond to the pilot's instruction because this command is based on an erroneous measurement of the pressure downstream of the high-pressure compressor. A command is therefore one or more instructions issued by the ECU to indicate to the engine components how they should operate or modify their operating mode to comply with the pilot's instruction.

[0013] The terms "pumping protection limit" and "extinguishing protection limit" refer to two maximum and minimum fuel flow rate limits for the engine, designed to prevent pumping or engine extinguishing. These values ​​replace the limits normally calculated by the engine control unit, as the latter are incorrect due to the pressure measurement anomaly.

[0014] The term "low power shutdown limit" refers to a limit value that prevents the motor from shutting down when it is in a low power operating mode. This value ensures that the motor will not shut down even if it enters a low power operating mode and despite a pressure measurement failure.

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

[0016] The measurement failure is detected by the determination of two indicators: a significant difference between the measured pressure and the pressure model, this model allowing the theoretical evaluation of the temperature at the outlet of the HP compressor; and an inconsistency between the behavior of the engine with the setpoint sent to it by the computer, which is determined according to the erroneous pressure measurement.

[0017] Compensation is then implemented to compensate for the failure of the pressure measurement. This compensation includes replacing the pressure measurement with the theoretical pressure calculated by the pressure model, which ensures that the engine control unit (ECU) constructs its engine control commands in accordance with the actual pressure at the outlet of the high-pressure compressor, i.e., the actual pressure downstream of said compressor and not the abnormal pressure. The compensation also includes determining three stops to replace those conventionally calculated by the The engine control unit (ECU) is used to define the engine's operating limits. The normally calculated limits are incorrect here because the pressure measurement does not match the actual pressure inside the engine. The compensation system therefore replaces these values ​​with values ​​that ensure the engine operates without major malfunctions, specifically without the risk of surging or stalling.

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

[0019] Thus, the pilot retains control of the aircraft even if there is a failure in the pressure measurement at the outlet of the high-pressure compressor. Compensation can continue until the failure is resolved, for example by a capillary protection device, or until the aircraft is shut down once its flight is complete.

[0020] Furthermore, the proposed method avoids sensitivity to a common-mode failure. Indeed, a redundant pressure measurement system can be implemented and still fail because the pressure is carried by the same capillary tube, which is duplicated at the redundant computers.

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

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

[0023] In one embodiment, the method according to the invention comprises: • Modify the engine speed based on the pressure obtained from the pressure model, the pumping protection limit, the extinction protection limit, and the low power extinction limit.

[0024] The engine operating regime is then corrected to conform to 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 value measured by the theoretical value.

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

[0028] In other words, the engine operating mode inconsistency is detected when: • Although the computer instructs the engine to reach maximum fuel flow, meaning that the engine speed is expected to increase, the engine speed remains unchanged or decelerates; or • Although the computer instructs the engine to reach the minimum fuel flow rate, meaning that the engine speed is expected to decelerate, the engine speed actually accelerates.

[0029] A second aspect of the invention relates to a device configured to implement the process 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, lead the latter to implement the steps of the process 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, lead the computer to carry out the steps of the process 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 illustrative purposes only and are in no way limiting of the invention. • Fig. 1 is a synoptic diagram illustrating the sequence of steps of a process for compensating for a failure in measuring pressure in an aircraft engine, according to one embodiment. • Fig. 2 is a schematic representation of a device configured to implement a method for compensating for a pressure measurement failure in an aircraft engine, according to one embodiment. • Fig. 3 is a diagram illustrating the determination of a pressure estimate by a theoretical model according to one embodiment. • Fig. 4 is a synoptic diagram illustrating the implementation according to one embodiment of the process of compensating for a failure in measuring pressure in an aircraft engine. DETAILED DESCRIPTION

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

[0036] As described below, the invention proposes a method for compensating for a pressure measurement failure in an aircraft engine. This method allows erroneous data processed by a computer to be replaced with corrected or alternative data. The computer, also known as a FADEC (Full Authority Digital Engine Control), is responsible for translating the pilot's flight instructions into commands for the aircraft engine components. The method comprises two main steps: the first detects the pressure measurement failure, and the second compensates for the failure.

[0037] The sequence of steps of process 100 is illustrated in [Fig.1].

[0038] Step 110, which involves detecting a pressure measurement failure, is implemented first. This detection is based on evaluating two indicators representative of this failure: a comparison of the measured pressure to a theoretical model, and the identification of inconsistencies between the actual operation of the engine and the setpoint issued by the computer.

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

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

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

[0042] The pressure model is a predefined theoretical model, for example, a model known in itself or a model established by an operator. The model can therefore be any model that allows the estimation of a theoretical pressure that the computer should detect via its sensing 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 exceeds 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 engine components, and / or their professional expertise.

[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, a measurement anomaly is detected when the difference between the measured value and the theoretical value exceeds a predefined threshold. The difference between these two values ​​is, for example, equal to the absolute value of the difference between them. Other types of difference can be used, such as the absolute value of the difference of the squares of the values, the sum of the values, the sum of the squares of the values, etc.

[0045] Step 110 also includes a step 112 for detecting inconsistencies between the actual motor operation and the setpoint issued by the computer. In particular, this step serves to detect inconsistencies between the motor's operating speed and the setpoint. The motor's operating speed here refers to the rotational speed of shafts N1 and / or N2.

[0046] An inconsistency can be detected when the engine speed decelerates or remains steady while the setpoint indicates that the engine speed should increase. Alternatively, an inconsistency can be detected when the engine speed increases 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 pilot's instruction. The fuel flow rate indicated by the instruction and determined by the computer is bounded by a maximum threshold and a minimum threshold, preventing the engine from entering pumping or stalling. Consequently, when the fuel flow rate indicated in the instruction is equal to the maximum or minimum threshold, the engine speed is expected to increase or decrease, respectively.

[0048] Therefore, the inconsistency between the engine's operating mode and the setpoint is detected when: • The engine operating mode corresponds to deceleration or a stationary speed while the engine fuel flow rate is equal to the maximum threshold; or • The engine operating regime 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 inconsistency of the operating regime are detected, then the measurement failure is detected.

[0051] Step 120, which compensates for the failure, is then implemented. This step comprises three stages, described below, relating to the substitution of the data required by the computer to determine the setpoints with 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 then use a pressure value representative of the actual pressure at the outlet of the high-pressure compressor to construct the setpoints, in particular to indicate the fuel flow rate 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 shutdown protection limit. These limits serve to condition the control unit so that the fuel flow rate it calculates and indicates in its setpoint does not exceed the pumping protection limit and is not less than the shutdown protection limit. When there is no fault, these limits are calculated by the control unit using known techniques. However, in the event of a fault, the conventional approaches fail and lead to the inconsistency mentioned above. Consequently, the two determined limits are substituted for the conventionally determined limits at the control unit level. The control unit therefore takes these determined limits into account, and no longer the conventionally determined limits, to construct its setpoint and calculate the corresponding fuel flow rate.

[0054] The pumping protection and extinction 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 NI or shaft N2.

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

[0056] Step 120 finally includes a step 123 for determining a low-power shutdown limit. This limit indicates a value below which the fuel flow rate must not fall to avoid the risk of the engine shutting down while 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 corresponding fuel flow rate.

[0057] The low-power extinction limit is determined from a body regime The high pressure of the engine, i.e., the rotational speed of the engine itself, is determined by the engine's rotational speed. The low-power shutdown limit can be calculated, either independently or jointly, from a predefined limit of the high-pressure compressor outlet pressure. This predefined limit is defined by the pressure model and / or by the operator based on the aircraft type and model, the characteristics of the engine control unit and / or engine components, and / or their professional expertise.

[0058] The process 100 may also include a step 130 of modifying the engine operating speed. 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 motor speed is therefore modified from the pressure obtained by the pressure model, the pumping protection limit, the extinction protection limit and the low power extinction limit.

[0060] The method 100 can be implemented by a device for compensating for a pressure measurement failure in an aircraft engine. This device is therefore configured to implement the method 100. Such a device 10 comprises, for example, and as schematically shown in [Fig. 2], a circuit including a processor 11 and volatile or non-volatile memory 12. The memory 12 is capable of storing instructions which, when implemented by the processor 11, cause the processor 11 to carry out the steps of the method 100. The device 10 is, for example, a computer.

[0061] The circuit can alternatively include an electronic board whose steps of the process of the invention are described in silicon, or a programmable electronic chip such as an FPGA chip (for "Field-Programmable Gate Array" in English).

[0062] The device 10 also includes 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 specifically, the input interface 13 is capable of receiving the measured pressure value. The input interface 12 is also capable of receiving the setpoint issued by the computer and the operating speed of the engine. In particular, the input interface 12 is capable of receiving data indicating the operating speed of said engine.

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

[0064] The device 10 can, moreover, be configured to perform calculations of the pressure model used by the process 100 and mentioned above. The device 10 Therefore, it includes, in its memory 12, instructions which, when implemented by the processor 11, lead the latter to perform pressure model calculations. Alternatively, the pressure model calculations are performed by an auxiliary module dedicated to these calculations. This module is thus configured to provide the device 10 with the pressure estimated by the model, that is, 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 an annex 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 can be a single device 10 for both computers or also be redundant, i.e. there is a device 10 for or in each computer.

[0067] By way of illustration, an example of implementation of process 100, in particular step 110, is detailed below in support of figures 3 and 4.

[0068] The pressure measurement is carried out using a capillary tube connected on one side to the outlet of the HP compressor 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 is estimated using the following pressure model. This model is an example of simplified pressure modeling in a turbomachine and has the advantage of being simple and quick to implement. This model is suitable for twin-spool, twin-flow turbojets.

[0071] As illustrated in [Fig. 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 for the low-pressure compressor speed, i.e., its rotational speed, expressed as a percentage of the maximum speed of the high-pressure compressor, and denoted PCNX2R. In the model, the low-pressure compressor speed, denoted X„12, is used to determine X„12R such that 127? = ^12 ' °ù -^12 is the estimated or measured temperature at the outlet of the low-pressure compressor. The value of PCNX2R is expressed as a function of X„12 / ?, for example as a ratio with the maximum speed of the high-pressure compressor, denoted Xn25MAX.

[0073] The second model M2 takes as input a value for the high-pressure compressor speed, that is, its rotational speed, expressed as a percentage of the maximum speed of this high-pressure compressor, and denoted PCN25R. In the model, the low-pressure compressor speed, denoted Xn25, is used to determine Xn25R such that Xn25R = 25, where T25 is the estimated or measured temperature at the outlet of the low-pressure compressor. The value of PCN25R is expressed as a function of Xn25R, for example, as a ratio with Xn25MAX.

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

[0075] The first pressure calculated by the first model M1 is multiplied by an estimate, or a measurement, of the total pressure PT2 upstream of the low-pressure compressor rotor. This multiplication produces an intermediate pressure value. This intermediate value is then multiplied by the second calculated pressure. The result of this second multiplication is the theoretical value PS3th of the pressure estimated by the pressure model.

[0076] The theoretical value PS3th is then compared, in step 111 and as illustrated in Figure 4, to the measured value, which is denoted PS3. The accuracy of this model is considered to be 30% and is denoted e30. The measurement anomaly is thus determined when: • PS3 £ PS3d 1 - g); or when • PS3> PS3th(l + e)-

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

[0078] In the example given, in step 112 and as illustrated in [Fig. 4], the inconsistency is detected when: • WF32c.,w / = WF32„U1X and that dNi < q ; or when said WF32^ = WF32„î / n and that £5 » n-dt '

[0079] The fuel flow rate indicated in the computer's setpoint is denoted WF32cmd. The maximum permitted fuel flow rate for the engine, also determined by the computer, is denoted WF3%nax. The minimum permitted fuel flow rate for the engine, also determined by the computer, is denoted WF32min. The time derivative of the rotational speed of shaft N2 is denoted , where r is time. dt

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

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

[0082] The measurement failure is then detected in two cases: • When the inconsistency arises because the N2 shaft speed indicates deceleration or stagnation, while the control fuel flow rate is equal to the maximum fuel flow rate and the measured pressure value is lower than the theoretical pressure value minus the error; or • When the inconsistency arises from 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 motor is not stalled, that is to say that the motor is not pumping, and provided that said motor 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 extinguishing protection limit are calculated and replace the maximum and minimum fuel flow rates conventionally calculated by the computer such as: = ( 1 + e'i.ClP,,,.,., PS^nsjb ; "WF32 ml „=(le).ClP m "PS3^T25 / b■

[0086] The value of the coefficient b is determined by the operator and is, for example, b = 288.15. C / Pma* is an upper limit that restricts the fuel flow to protect the high-pressure compressor from pumping due to excessive fuel consumption during pumping related to thermal vortexing caused by the fuel level. The value of C / Pmax can be determined by any known prior art method.

[0087] In addition, control loops are implemented to limit the acceleration and deceleration rates of the HP body. The acceleration rate is thus limited by a maximum value denoted dN2 and a minimum value denoted ^3. These dl max dt min Regulation loops, known in themselves, allow for the application of more restrictive protection but nevertheless allow for acceleration and deceleration in the event of a pressure measurement failure.

[0088] Consequently, the engine is doubly protected from pumping or extinction by the replacement of 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, and then send to the devices a new instruction based on this data to respond to the piloting instruction required by the pilot.

Claims

Demands

1. A method (100) for compensating for a pressure measurement failure in an aircraft engine, the method comprising: - Detecting (110) a pressure measurement failure by: • 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 • Detecting (112) an inconsistency in the operating regime of the aircraft engine with a setpoint; - Compensating (120) for the measurement failure by: • Substituting (121) the pressure measurement with a pressure obtained from the pressure model; • 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;• Determination (123) of a low power extinction 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 extinction limit being added to the pumping and extinction protection limits.

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

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

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

5. A method according to any one of the preceding claims, wherein the setpoint is a fuel flow rate and wherein the inconsistency is detected when: - The engine operating speed corresponds to deceleration or a stationary speed while the fuel flow rate is equal to a maximum threshold; or - The engine operating speed corresponds to acceleration while the fuel flow rate is equal to a minimum threshold.

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

7.

8. Aircraft comprising the device (10) according to the preceding claim. Computer program product comprising instructions which, when the program is executed on a computer, cause the computer to carry out the steps of the process (100) according to any one of claims 1 to "

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

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