Methods for predicting fault signals based on altitude and Mach number in aero-engines

Abstract

A method for predicting altitude and Mach number fault signals in an aero-engine is disclosed. The method involves the aircraft transmitting three signals to the aero-engine: altitude, Mach number, and whether the data is valid. The validity of the data is determined. If valid, the engine updates its recorded altitude, Mach number, and static temperature drift for future signal transmission. If invalid, the partial derivatives of theoretical total temperature and theoretical total pressure with respect to altitude and Mach number are calculated using the previously recorded altitude, Mach number, and static temperature drift. The partial derivatives are used to predict the current altitude and Mach number, and the engine updates its recorded altitude and Mach number for future signal transmission. The accuracy is verified through simulation using a component-level digital model of the aero-engine.

Description

Technical Field

[0001] This invention belongs to the field of aerospace, and in particular, it is a method for predicting fault signals of altitude and Mach number in aero-engines. Background Technology

[0002] The performance tuning of aircraft engines relies on precise flight data, including altitude and Mach number signals. These signals are typically provided by the aircraft's sensor systems and transmitted in real time to the engine control unit to achieve optimal engine performance and fuel efficiency. However, in highly dynamic flight environments, such as high-speed or atmospheric edge operations, sensor systems may experience brief malfunctions, leading to interruptions in altitude and Mach number signals. Such signal interruptions can affect engine performance tuning and even threaten flight safety.

[0003] Existing technologies are generally unable to handle such signal loss situations, thus often failing to accurately reflect the actual flight status under highly dynamic conditions at all times, or to respond promptly to rapidly changing flight environments. Therefore, there is an urgent need to develop a reliable altitude and Mach number prediction method that can accurately predict the flight status using existing engine measurement parameters when aircraft sensor signals fail, ensuring that the engine control system can continue to make effective performance adjustments, and guaranteeing the normal operation and flight safety of the aircraft.

[0004] The information disclosed in the background section is only intended to enhance the understanding of the background of the present invention, and therefore may contain information that does not constitute prior art known to those skilled in the art. Summary of the Invention

[0005] To address the problems existing in the prior art, this invention proposes a method for predicting altitude and Mach number fault signals of aero-engines. When the aircraft sensor signals fail, the method can accurately predict the flight status using the engine's existing measurement parameters, ensuring that the engine control system can continue to perform effective performance adjustments.

[0006] The objective of this invention is achieved through the following technical solution: a method for predicting altitude and Mach number fault signals in an aero-engine includes:

[0007] Step S100: The aircraft transmits three signals to the aircraft engine: altitude, Mach number, and whether the data is valid;

[0008] Step S200: Determine if the data is valid;

[0009] Step S300: If the data is valid, the aircraft engine updates the recorded altitude, Mach number, and static temperature drift for use in the next signal transmission;

[0010] Step S400: If the data is invalid, calculate the partial derivatives of the theoretical total temperature and theoretical total pressure with respect to the height and Mach number using the previously recorded height, Mach number, and static temperature drift.

[0011] Step S500: Predict the current altitude and Mach number using partial derivatives, and update the current altitude and Mach number of the aero-engine for use in the next signal transmission;

[0012] Step S600: Perform simulation using the engine component-level digital model to verify accuracy.

[0013] In the aforementioned method for predicting altitude and Mach number fault signals of an aero-engine, in step S100, the aircraft's sensors detect the altitude and Mach number at intervals of steptime. The detected altitude and Mach number are transmitted to the aero-engine port. The three signals, altitude, Mach number, and whether the data is valid, are denoted as H0, M0, and OK, respectively. The OK signal is either 0 or 1.

[0014] In the aforementioned method for predicting altitude and Mach number fault signals of an aero-engine, in step S200, the aero-engine determines whether the data is valid (OK). If the value is 1, it means that the altitude and Mach number are valid and can be used subsequently. If the data is valid (OK) is not equal to 1, it means that the altitude and Mach number are invalid.

[0015] In the aforementioned method for predicting altitude and Mach number fault signals for an aero-engine, in step S300, if the data is valid, the data validity signal is equal to 1, and the recorded altitude H1, Mach number M1, and temperature drift value dt are updated.

[0016] The inertial filtering formula for temperature drift is as follows:

[0017] dt = dt + steptime * (T1 - T)

[0018] In the formula, steptime is the step time, the interval between signals transmitted from the aircraft to the aero-engine, dt on the right side of the equation is the previous inertia value, dt on the left side of the equation is the current filtered value, T1 is the total temperature measured by the engine, and T is the theoretical total temperature calculated from altitude and Mach number.

[0019] In the aforementioned method for predicting altitude and Mach number fault signals of an aero-engine, in step S400, when the data is invalid, the altitude H0 and Mach number M0 transmitted from the aircraft to the aero-engine are either nonexistent or inaccurate. The partial derivatives of the theoretical total temperature T and theoretical total pressure P with respect to altitude and Mach number are calculated using the recorded altitude values ​​H1 and Mach number M1. The calculation method is as follows:

[0020]

[0021]

[0022]

[0023]

[0024] In the formula, ΔH and ΔM are the small changes in height and Mach number, respectively;

[0025] The calculation methods for theoretical total temperature T and theoretical total pressure P are as follows:

[0026] T = T0 * (1 + 0.5(k-1)M1) 2 )

[0027]

[0028] In the formula, T0 and P0 are the theoretical ambient static temperature and static pressure, respectively, and k is the adiabatic coefficient.

[0029] By substituting altitude and Mach number into the interpolation function under standard day conditions, the theoretical static temperature T0 and theoretical static pressure P0 of the environment are obtained.

[0030] The formulas for calculating the theoretical static temperature T0 and the theoretical static pressure P0 are as follows:

[0031]

[0032]

[0033] In the formula: H[i] represents the i-th height interpolation point, and H is located between H[i-1] and H[i]; T[i] represents the i-th ambient static temperature interpolation point; P[i] represents the i-th theoretical ambient static pressure interpolation point; dt represents the static temperature drift.

[0034] In the aforementioned method for predicting altitude and Mach number fault signals of an aero-engine, step S500 utilizes the partial derivatives of theoretical total temperature T and theoretical total pressure P with respect to altitude and Mach number. The formulas for calculating the changes in height and Mach number relative to the previous value are as follows:

[0035]

[0036]

[0037] In the formula, T1 is the total temperature measured by the engine in this test, and P1 is the total pressure measured by the engine in this test; T is the theoretical total temperature, and P is the theoretical total pressure.

[0038] The predicted height and Mach number are obtained and recorded as update values. The calculation formula is as follows:

[0039] H1=H1+dH

[0040] M1 = M1 + dM.

[0041] In the aforementioned method for predicting altitude and Mach number fault signals in an aero-engine, step S600 involves using an aero-engine component-level digital model to set random altitude and Mach number variations within a time step, and setting temperature drift to change linearly with time. The input parameters are the output parameters of the aero-engine component-level digital model each time, including altitude, Mach number, total temperature, total pressure, and whether the data is valid. The output parameters are the predicted altitude and Mach number. The altitude and Mach number parameters output by the aero-engine component-level digital model are considered as the actual data transmitted from the aircraft to the aero-engine. A fault time period is set at different altitudes, i.e., OK=0, and OK=1 for the rest of the time. Finally, the test results are statistically analyzed, recording the actual altitude, predicted altitude, and absolute altitude error of the engine model at the point of maximum cumulative error within the fault time; the actual Mach number, predicted Mach number, and absolute Mach number error of the engine model.

[0042] The maximum error between the predicted and actual values ​​is calculated using the following formula based on different fault times:

[0043] ε H =Max|H1-H * |

[0044] ε M =Max|M1-M * |

[0045] In the formula, H1 and M1 are the predicted values ​​obtained using the prediction method, and H... * M * The state values ​​output for the digital model of aero-engine components.

[0046] Compared with existing technologies, this invention has the following advantages: The total temperature and total pressure parameters in an aero-engine are measured by the aero-engine itself and are not susceptible to failure; the altitude and Mach number signal is transmitted from the aircraft and is prone to failure. Therefore, during temperature drift recording, although the aircraft only transmits the altitude and Mach number, the temperature drift can be calculated using formulas because the total temperature and total pressure measured by the engine are accurate. This invention can automatically switch to predictive mode during flight, improving flight safety and reducing the risk of flight accidents caused by the loss of altitude and Mach number data. This invention takes into account the drift of ambient static temperature, making it applicable to both standard and non-standard weather conditions. This invention considers the altitude and Mach number at the previous moment, acting as an inertial filter to prevent sudden changes in altitude and Mach number data during faults.

[0047] Based on partial derivatives and engine models, this invention offers high accuracy and real-time performance, providing pilots with reliable altitude and Mach number information within a short time. The algorithm is relatively simple to implement and easily integrated into existing aero-engine control systems, requiring no additional hardware and reducing system complexity and cost. Attached Figure Description

[0048] Various other advantages and benefits of the present invention will become apparent to those skilled in the art upon reading the detailed description of the preferred embodiments below. The accompanying drawings are for illustrative purposes only and are not intended to limit the invention. It is obvious that the drawings described below are merely some embodiments of the invention, and those skilled in the art can obtain other drawings based on these drawings without any inventive effort. Furthermore, the same reference numerals denote the same parts throughout the drawings.

[0049] In the attached diagram:

[0050] Figure 1 This is a flowchart illustrating a method for predicting altitude and Mach number fault signals for an aero-engine, provided in one embodiment of this disclosure.

[0051] The present invention will be further explained below with reference to the accompanying drawings and embodiments. Detailed Implementation

[0052] Specific embodiments of the invention will now be described in more detail with reference to the accompanying drawings. While specific embodiments of the invention are shown in the drawings, it should be understood that the invention can be implemented in various forms and should not be limited to the embodiments set forth herein. Rather, these embodiments are provided to enable a more thorough understanding of the invention and to fully convey the scope of the invention to those skilled in the art.

[0053] It should be noted that certain terms are used in the specification and claims to refer to specific components. Those skilled in the art will understand that different terms may be used to refer to the same component. This specification and claims do not distinguish components based on differences in terminology, but rather on differences in function. The terms "comprising" or "including" used throughout the specification and claims are open-ended and should be interpreted as "comprising but not limited to." The following descriptions are preferred embodiments for carrying out the invention; however, these descriptions are for the purpose of understanding the general principles of the specification and are not intended to limit the scope of the invention. The scope of protection of this invention is determined by the appended claims.

[0054] To facilitate understanding of the embodiments of the present invention, further explanations and descriptions will be provided below with reference to the accompanying drawings and specific embodiments. The accompanying drawings do not constitute a limitation on the embodiments of the present invention.

[0055] like Figure 1 As shown, the methods for predicting fault signals related to altitude and Mach number in aero-engines include:

[0056] Step S100: The aircraft transmits three signals to the aircraft engine: altitude, Mach number, and whether the data is valid;

[0057] Step S200: Determine if the data is valid;

[0058] Step S300: If the data is valid, the aircraft engine updates the recorded altitude, Mach number, and static temperature drift for use in the next signal transmission;

[0059] Step S400: If the data is invalid, calculate the partial derivatives of the theoretical total temperature and theoretical total pressure with respect to the height and Mach number using the previously recorded height, Mach number, and static temperature drift.

[0060] Step S500: Predict the current altitude and Mach number using partial derivatives, and update the current altitude and Mach number of the aero-engine for use in the next signal transmission;

[0061] Step S600: Perform simulation using the engine component-level digital model to verify accuracy.

[0062] In a preferred embodiment of the method for predicting altitude and Mach number fault signals of an aero-engine, in step S100, the aircraft's sensors detect the altitude and Mach number every steptime interval. The detected altitude and Mach number are transmitted to the port of the aero-engine. The three signals, altitude, Mach number, and whether the data is valid, are denoted as H0, M0, and OK, respectively. The OK signal is either 0 or 1.

[0063] In a preferred embodiment of the method for predicting altitude and Mach number fault signals of an aero-engine, in step S200, the aero-engine determines whether the data is valid (OK). If the value is 1, it means that the altitude and Mach number are valid and can be used later. If the data is valid (OK) is not equal to 1, it means that the altitude and Mach number are invalid.

[0064] In a preferred embodiment of the method for predicting altitude and Mach number fault signals in an aero-engine, in step S300, if the data is valid, the data validity signal is equal to 1, and the recorded altitude H1, Mach number M1, and temperature drift value dt are updated.

[0065] The inertial filtering formula for temperature drift is as follows:

[0066] dt = dt + steptime * (T1 - T)

[0067] In the formula, steptime is the step time, the interval between signals transmitted from the aircraft to the aero-engine, dt on the right side of the equation is the previous inertia value, dt on the left side of the equation is the current filtered value, T1 is the total temperature measured by the engine, and T is the theoretical total temperature calculated from altitude and Mach number.

[0068] In a preferred embodiment of the method for predicting altitude and Mach number fault signals of an aero-engine, in step S400, when the data is invalid, the altitude H0 and Mach number M0 transmitted from the aircraft to the aero-engine are nonexistent or inaccurate. The partial derivatives of the theoretical total temperature T and theoretical total pressure P with respect to altitude and Mach number are calculated using the recorded altitude values ​​H1 and Mach number M1. The calculation method is as follows:

[0069]

[0070]

[0071]

[0072]

[0073] In the formula, ΔH and ΔM are the small changes in height and Mach number, respectively;

[0074] The calculation methods for theoretical total temperature T and theoretical total pressure P are as follows:

[0075] T = T0 * (1 + 0.5(k-1)M1) 2 )

[0076]

[0077] In the formula, T0 and P0 are the theoretical ambient static temperature and static pressure, respectively, and k is the adiabatic coefficient.

[0078] By substituting altitude and Mach number into the interpolation function under standard day conditions, the theoretical static temperature T0 and theoretical static pressure P0 of the environment are obtained.

[0079] The formulas for calculating the theoretical static temperature T0 and the theoretical static pressure P0 are as follows:

[0080]

[0081]

[0082] In the formula: H[i] represents the i-th height interpolation point, and H is located between H[i-1] and H[i]; T[i] represents the i-th ambient static temperature interpolation point; P[i] represents the i-th theoretical ambient static pressure interpolation point; dt represents the static temperature drift.

[0083] In a preferred embodiment of the method for predicting altitude and Mach number fault signals of an aero-engine, in step S500, the partial derivatives of theoretical total temperature T and theoretical total pressure P with respect to altitude and Mach number are used. The formulas for calculating the changes in height and Mach number relative to the previous value are as follows:

[0084]

[0085]

[0086] In the formula, T1 is the total temperature measured by the engine in this test, and P1 is the total pressure measured by the engine in this test; T is the theoretical total temperature, and P is the theoretical total pressure.

[0087] The predicted height and Mach number are obtained and recorded as update values. The calculation formula is as follows:

[0088] H1=H1+dH

[0089] M1 = M1 + dM.

[0090] In a preferred embodiment of the method for predicting altitude and Mach number fault signals in an aero-engine, in step S600, a component-level digital model of the aero-engine is used to set random altitude and Mach number variations within a step time, and to set temperature drift to change linearly with time. The input quantities are the parameters output by the component-level digital model of the aero-engine each time, including altitude, Mach number, total temperature, total pressure, and whether the data is valid. The output quantities are the predicted altitude and Mach number. The altitude and Mach number parameters output by the component-level digital model of the aero-engine are regarded as the real data transmitted from the aircraft to the aero-engine. A fault time period is set at different altitudes, i.e., OK=0, and OK=1 for the rest of the time. Finally, the test results are statistically analyzed, and the actual altitude, predicted altitude, and absolute altitude error of the engine model at the point of maximum cumulative error within the fault time are recorded; the actual Mach number, predicted Mach number, and absolute Mach number error of the engine model are also recorded.

[0091] The maximum error between the predicted and actual values ​​is calculated using the following formula based on different fault times:

[0092] ε H =Max|H1-H * |

[0093] ε M =Max|M1-M * |

[0094] In the formula, H1 and M1 are the predicted values ​​obtained using the prediction method, and H * M * State values ​​output for aero-engine component-level digital models

[0095] In a preferred embodiment of the method for predicting altitude and Mach number fault signals of an aero-engine, in step S400, the small change in altitude is set to 50 meters, the small change in Mach number is set to 0.1, and the adiabatic coefficient k is set to 1.4. In step S600, the random altitude change and Mach number change within a step time are set, and the accuracy is verified using an engine model. The random altitude range is 0-3.5m, the Mach number change range is 0-0.00023, and the temperature drift change is fixed at 0.0025K. The iteration is performed 6000 times. At altitudes of 1km, 2km, 5km, 10km, and 14km, the engine is stopped for 0.5s. At these five altitudes, the data validity signal 0K is set to 0, and at all other times, the data validity signal OK is set to 1. Finally, the test results are statistically analyzed, and the actual altitude, predicted altitude, and absolute altitude error of the engine model at the point of maximum cumulative error in the above five time periods are recorded; the actual Mach number, predicted Mach number, and absolute Mach number error of the engine model are also recorded.

[0096] In one embodiment, the method for predicting altitude and Mach number fault signals of an aero-engine includes...

[0097] S100: The aircraft transmits signals to the aircraft engine indicating altitude, Mach number, and whether the data is valid.

[0098] In this step, to minimize redundancy in characterizing the state information of the engine under test, only four parameters—height H0, Mach number M0, total temperature T1, and total pressure P1—are selected as the operating data to characterize the engine. The engine model is run to the height H0 and Mach number M0, and the above four state signals, along with the data validity signal OK, are set as input quantities and fed into the prediction function.

[0099] S200: Determine if the data is valid;

[0100] In this step, OK is checked every so often. When OK equals 1, the height Mach number signal is considered to be without fault and there is no need to predict the height Mach number. When OK equals 0, the height Mach number signal is considered to be without fault and the predicted height Mach number needs to be used instead.

[0101] S300: If the data is valid, the aircraft engine updates the recorded altitude, Mach number, and static temperature drift for use in the next signal transmission.

[0102] In this step, the variables H1 and M1, used to record height and Mach number, are assigned values, as shown in the following formula:

[0103] H1 = H0

[0104] M1 = M0

[0105] The inertial filtering formula for temperature drift is as follows:

[0106] dt = dt + steptime * (T1 - T)

[0107] In the formula, steptime is the step time, set to 0.025s, which is the interval between the simulated aircraft transmitting signals to the engine. dt on the right side of the equation is the previous inertia value, and dt on the left side of the equation is the current filtered calculation value. T1 is the total temperature obtained from the engine model, and T is the theoretical total temperature calculated from the altitude Mach number, as explained in step S400.

[0108] S400: If the data is invalid, calculate the partial derivatives of the theoretical total temperature and theoretical total pressure with respect to the height and Mach number using the previously recorded height, Mach number, and static temperature drift.

[0109] In this step, some data points are selected where OK = 0, indicating that the current data is invalid, meaning that the altitude H0 and Mach number M0 transmitted from the aircraft to the aero-engine are nonexistent or inaccurate. Using the recorded altitude values ​​H1 and Mach number M1, the partial derivatives of the theoretical total temperature and theoretical total pressure with respect to altitude and Mach number are calculated. The calculation method is as follows:

[0110]

[0111]

[0112]

[0113]

[0114] In the formula, ΔH and ΔM are the small changes in Mach number at height. Let ΔH = 50m and ΔM = 0.1. In one embodiment, the small change in height is taken as 50, the small change in Mach number is taken as 0.1, and the adiabatic coefficient k is taken as 1.4.

[0115] The calculation methods for theoretical total temperature T and theoretical total pressure P are as follows:

[0116] T = T0 * (1 + 0.5(k-1)M1) 2 )

[0117]

[0118] In the formula, T0 and P0 are the theoretical static temperature and theoretical static pressure, respectively, and k is the adiabatic coefficient, which is taken as 1.4.

[0119] By substituting the high Mach number into the interpolation function under standard day conditions, the theoretical ambient static temperature T0 and theoretical ambient static pressure P0 are obtained.

[0120] The formulas for calculating the theoretical static temperature T0 and the theoretical static pressure P0 are as follows:

[0121]

[0122]

[0123] In the formula: H[i] represents the i-th height interpolation point, and H is located between H[i-1] and H[i]; T[i] represents the i-th theoretical environmental static temperature interpolation point; P[i] represents the i-th theoretical environmental static pressure interpolation point; dt represents the static temperature drift.

[0124] S500: The partial derivative is used to predict the current altitude and Mach number, and the aero engine updates the current altitude and Mach number for use in the next signal transmission.

[0125] In this step, the partial derivatives obtained in step S400 are used to calculate the changes in height and Mach number relative to the previous result. The calculation formula is as follows:

[0126]

[0127]

[0128] In the formula, the partial derivatives The calculation is shown in step S400; T1 is the total temperature output by the engine model, P1 is the total pressure output by the engine model; T is the theoretical total temperature obtained from the previous altitude Mach number calculation, and P is the theoretical total pressure obtained from the previous altitude Mach number calculation. As shown in step S400.

[0129] The predicted height and Mach number are obtained and recorded as update values. The calculation formula is as follows:

[0130] H1=H1+dH

[0131] M1 = M1 + dM.

[0132] S600: Simulation using an engine model to verify accuracy.

[0133] In this step, a component-level digital model of a certain engine type is used, with a steptime of 0.025s. Random altitude and random Mach number are varied each time, with the random altitude ranging from 0 to 3.5m, the varying Mach number ranging from 0 to 0.00023, and the temperature drift fixed at 0.0025K. This process is repeated 6000 times. At approximately altitudes of 1km, 2km, 5km, 10km, and 14km, the engine is stopped for 0.5s, with OK set to 0 for 0.5s. At all other times, OK is set to 1. For the prediction method, the inputs are the parameters output by the component-level digital model of the aero-engine, including altitude, Mach number, total temperature, total pressure, and whether the data is a valid signal. The outputs are the predicted altitude and Mach number. The altitude and Mach number parameters output by the component-level digital model of the aero-engine are considered as the actual data transmitted from the aircraft to the aero-engine. Finally, the test results were statistically analyzed, and the actual height, predicted height and absolute height error of the engine model at the point of maximum cumulative error were recorded for the above five time periods; the actual Mach number, predicted Mach number and absolute Mach number error of the engine model were also recorded.

[0134] The maximum error between the predicted and actual values ​​is calculated using the following formula based on different fault times:

[0135] ε H =Max|H1-H * |

[0136] ε M =Max|M1-M * |

[0137] In the formula, H1 and M1 are the predicted values ​​obtained using the prediction method, and H * M * The state values ​​output for the digital model of aero-engine components.

[0138] Table 1 Height Prediction Test

[0139]

[0140] Table 2 Mach number prediction test

[0141]

[0142] Simulation results are shown in Tables 1 and 2. As can be seen from the tables, the results obtained using the high Mach number fault signal prediction method are basically consistent with the actual results, demonstrating the effectiveness of the high Mach number fault signal prediction method disclosed herein.

[0143] Although embodiments of the present invention have been described above in conjunction with the accompanying drawings, the present invention is not limited to the specific embodiments and application fields described above. The specific embodiments described above are merely illustrative and instructive, and not restrictive. Those skilled in the art can make many other forms based on the guidance of this specification and without departing from the scope of protection of the claims of the present invention, and all of these are within the scope of protection of the present invention.

Claims

1. A method for predicting altitude and Mach number fault signals in an aero-engine, characterized in that, It includes the following steps, Step S100: The aircraft transmits three signals to the aircraft engine: altitude, Mach number, and whether the data is valid; Step S200: Determine if the data is valid; Step S300: If the data is valid, the aircraft engine updates the recorded altitude, Mach number, and static temperature drift for use in the next signal transmission; Step S400: If the data is invalid, calculate the partial derivatives of the theoretical total temperature and theoretical total pressure with respect to the height and Mach number using the previously recorded height, Mach number, and static temperature drift. Step S500: Use partial derivatives to predict the current altitude and Mach number, and update the recorded altitude and Mach number of the aero engine for use in the next signal transmission; Step S600: Perform simulation using the engine component-level digital model to verify accuracy; In step S400, when the data is invalid, the altitude transmitted by the aircraft to the aero-engine is... ,Mach number It does not exist or is not real, based on the recorded height value. ,Mach number Calculate the theoretical total temperature Theoretical total pressure Partial derivatives with respect to altitude and Mach number The calculation method is as follows: ， ， ， ， In the formula, , These are the minute changes in altitude and Mach number, respectively. Theoretical total temperature Theoretical total pressure The calculation method is as follows: ， ， In the formula, , These are the theoretical static temperature and the theoretical static pressure, respectively. The adiabatic coefficient is... By substituting altitude and Mach number into the interpolation function under standard day conditions, the theoretical ambient static temperature is obtained. static pressure relative to theoretical environment , Theoretical environment static temperature static pressure relative to theoretical environment The calculation formula is as follows: ， ， In the formula: Represents the i-th height interpolation point. lie in and between; This represents the i-th ambient static temperature interpolation point; This represents the i-th theoretical environmental static pressure interpolation point; This represents static temperature drift.

2. The method for predicting altitude and Mach number fault signals of an aero-engine according to claim 1, characterized in that, In step S100, at intervals of time... The aircraft's sensors detect altitude and Mach number, and the detected altitude and Mach number are transmitted to the engine ports. The altitude, Mach number, and data validity are recorded as three signals. , OK, the OK signal is either 0 or 1.

3. The method for predicting altitude and Mach number fault signals of an aero-engine according to claim 1, characterized in that, In step S200, the aircraft engine determines whether the data is valid (OK). If it is 1, it means that the altitude and Mach number are valid and can be used later. If the data is valid (OK) is not equal to 1, it means that the altitude and Mach number are invalid.

4. The method for predicting altitude and Mach number fault signals of an aero-engine according to claim 3, characterized in that, In step S300, if the data is valid, the data validity signal equals 1, and the record height is updated. ,Mach number and temperature drift value , The inertial filtering formula for temperature drift is as follows: ， In the formula The step time is the interval between signals transmitted from the aircraft to the aero-engine, and the right-hand side of the equation is... The previous inertia value is given on the left side of the equation. This is the value calculated for this filtering process. The total temperature measured for the engine. The theoretical total temperature is calculated based on altitude and Mach number.

5. The method for predicting altitude and Mach number fault signals of an aero-engine according to claim 1, characterized in that, In step S600, using an aero-engine component-level digital model, random altitude and Mach number variations are set within a step time, and temperature drift is set to change linearly with time. The inputs are the parameters output by the aero-engine component-level digital model each time, including altitude, Mach number, total temperature, total pressure, and whether the data is valid. The outputs are the predicted altitude and Mach number. The altitude and Mach number parameters output by the aero-engine component-level digital model are considered as the actual data transmitted from the aircraft to the aero-engine. A fault time is set at different altitudes, i.e., OK=0, and OK=1 at other times. Finally, the test results are statistically analyzed, and the actual altitude, predicted altitude, and absolute altitude error of the engine model at the point of maximum cumulative error within the fault time are recorded; the actual Mach number, predicted Mach number, and absolute Mach number error of the engine model are also recorded. The maximum error between the predicted and actual values ​​within different fault times is calculated using the following formula: ， ， In the formula, , These are the predicted values ​​obtained using the prediction method. , The state values ​​output for the digital model of aero-engine components.

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