Aeroengine fault monitoring system and control method thereof

Abstract

This disclosure relates to an aircraft engine fault monitoring system and its control method. The aircraft engine fault monitoring system includes: an engine body; and one or more laser positioning radars disposed around the engine body, configured to detect the size, position, and motion parameters of foreign objects. By distributing laser positioning radars around the engine body, the laser positioning radars can accurately detect the size, position, and motion parameters of foreign objects, significantly reducing negligence and misjudgment due to manual monitoring of parameters, and improving the accuracy of engine fault detection.

Description

Technical Field

[0001] This disclosure relates to an aircraft engine fault monitoring system and its control method. Background Technology

[0002] The statements herein are provided only as background information in connection with this disclosure and do not necessarily constitute prior art.

[0003] In an aero-engine (including turbofan engines, turboshaft engines, and turboprop engines), the airflow flows in the axial direction, passing sequentially through the compressor components that compress the airflow, the combustion chamber that heats the airflow, the turbine components that drive the compressor, and the nozzle that accelerates and ejects the high-temperature, high-pressure airflow.

[0004] During engine operation, foreign objects such as birds, sand, or hail may be drawn into the engine inlet and collide with the high-speed rotating impeller, potentially causing blade breakage, friction-induced fire, or other safety accidents. Unexpected foreign object ingestion must be strictly avoided during engine testing.

[0005] Currently, the diagnosis of foreign object ingestion, structural failure, and surge monitoring faults during aero-engine operation tests mainly relies on manual inspection of vibration sensors. This method can only detect vibration faults but cannot determine the cause of the fault or prevent its occurrence during the detection process. Summary of the Invention

[0006] One of the technical problems to be solved by this disclosure is to provide an aircraft engine fault monitoring system and its control method, which can improve the accuracy of engine fault detection.

[0007] Some embodiments of this disclosure provide an aircraft engine fault monitoring system, including: an engine body; and one or more lidar positioning radars disposed around the engine body, configured to detect the size parameters, position parameters, and motion parameters of foreign objects.

[0008] In some embodiments, foreign objects include air intake foreign objects, accessory splash foreign objects, and exhaust splash foreign objects.

[0009] In some embodiments, multiple laser positioning radars are distributed in different locations around the engine body to achieve redundant detection.

[0010] In some embodiments, the laser positioning radar includes a first laser probe and a second laser probe, the first laser probe being configured to detect and record the angular coordinates α1 and β1 of the foreign object relative to the first laser probe, and the second laser probe being configured to detect and record the angular coordinates α2 and β2 of the foreign object relative to the second laser probe.

[0011] In some embodiments, one or more acoustic wave collectors are also included, disposed around the engine body, and configured to collect acoustic wave signals from the engine body.

[0012] In some embodiments, multiple acoustic wave collectors are distributed in different locations around the engine body to achieve redundancy detection.

[0013] In some embodiments, at least one acoustic wave collector is positioned at a 60° angle to the engine axis of the engine body.

[0014] In some embodiments, the system further includes a digital controller and a host computer, the host computer being configured to receive signals from the laser positioning radar and the acoustic wave collector and to process the data, and the digital controller being configured to control the fuel supply and adjustable actuation mechanism of the engine body according to the data processed by the host computer, so as to realize engine speed control and stopping operation.

[0015] Some embodiments of this disclosure provide a control method for controlling the aforementioned aero-engine fault monitoring system, including:

[0016] When the laser positioning radar detects a foreign object in the inner air intake area, if the diameter of the foreign object is greater than the first preset diameter or its speed is greater than the first preset speed, the engine body is controlled to stop; otherwise, the engine body is controlled to reduce its speed to idle.

[0017] When the laser positioning radar detects a foreign object in the outer bypass air intake area, if the diameter of the foreign object is greater than the second preset diameter or its speed is greater than the second preset speed, the engine body is controlled to stop; otherwise, the engine body is controlled to reduce its speed to idle.

[0018] When the laser positioning radar detects a foreign object in the accessory area, if the diameter of the foreign object is greater than the third preset diameter or its speed is greater than the third preset speed, the engine body is controlled to stop; otherwise, the engine body is controlled to reduce its speed to slow.

[0019] When the laser positioning radar detects a foreign object in the exhaust zone, if the diameter of the foreign object is greater than a fourth preset diameter and its speed is greater than a fourth preset speed, the engine is stopped; otherwise, the engine speed is reduced to idle.

[0020] When the laser positioning radar detects a foreign object in the exhaust area of ​​the outer casing, if the diameter of the foreign object is greater than the fifth preset diameter and its speed is greater than the fifth preset speed, the engine body is controlled to stop; otherwise, the engine body is controlled to reduce its speed to idle.

[0021] Some embodiments of this disclosure provide a control method for controlling the aforementioned aero-engine fault monitoring system, including:

[0022] If the intensity of the sound wave collected by the sound wave collector is between a first preset intensity value and a second preset intensity value, the engine body is controlled to stop; and

[0023] If the intensity of the sound wave collected by the sound wave collector exceeds the second preset intensity value, the engine body will be stopped.

[0024] This disclosure improves engine fault detection accuracy by installing a laser positioning radar around the engine body, which can accurately detect the size, position, and motion parameters of foreign objects. This significantly reduces the negligence and misjudgment caused by manual monitoring of parameters. Attached Figure Description

[0025] To more clearly illustrate the technical solutions in the embodiments of this disclosure or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of this disclosure. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0026] Figure 1 This is a three-dimensional structural schematic diagram of some embodiments of the aircraft engine fault monitoring system disclosed herein;

[0027] Figure 2 These are the frequency domain intensity characteristics of the acoustic waves of an aero-engine at various speeds;

[0028] Figure 3 This is a diagram showing the frequency domain intensity characteristics of the sound waves of an aero-engine under normal operating conditions.

[0029] Figure 4 This is a diagram showing the intensity limits of acoustic waves from an aircraft engine.

[0030] Figure 5 It is a characteristic diagram of the total intensity of acoustic signals from an aero-engine;

[0031] Figure 6 This is a schematic diagram illustrating the effect of the angle and position of the acoustic sensor on the acoustic wave intensity.

[0032] Explanation of reference numerals in the attached figures

[0033] 1. Laser positioning radar; 2. Acoustic wave collector; 3. Digital controller; 4. Host computer; 5. Engine accessories; 6. Intake foreign objects; 7. Accessory splash foreign objects; 8. Exhaust splash foreign objects; 10. Engine body. Detailed Implementation

[0034] Various exemplary embodiments of the present disclosure will now be described in detail with reference to the accompanying drawings. The descriptions of the exemplary embodiments are merely illustrative and are in no way intended to limit the present disclosure or its application or use. The present disclosure may be implemented in many different forms and is not limited to the embodiments herein. These embodiments are provided so that the present disclosure will be thorough and complete, and will fully express the scope of the disclosure to those skilled in the art. It should be noted that, unless specifically stated otherwise, the relative arrangement of components and steps, the composition of materials, numerical expressions, and values ​​set forth in these embodiments should be interpreted as merely exemplary and not as limiting.

[0035] The terms "first," "second," and similar words used in this disclosure do not indicate any order, quantity, or importance, but are merely used to distinguish different parts. Words such as "including" or "contains" mean that the element preceding the word encompasses the element listed after it, and do not exclude the possibility of encompassing other elements as well. Terms such as "above," "below," "left," and "right" are used only to indicate relative positional relationships; when the absolute position of the described object changes, this relative positional relationship may also change accordingly.

[0036] In this disclosure, when a specific device is described as being located between a first device and a second device, an intermediary device may or may not be present between the specific device and the first or second device. When a specific device is described as being connected to other devices, the specific device may be directly connected to the other devices without an intermediary device, or it may be not directly connected to the other devices but have an intermediary device.

[0037] All terms used in this disclosure have the same meaning as understood by one of ordinary skill in the art to which this disclosure pertains, unless otherwise specifically defined. It should also be understood that terms defined in general dictionaries should be interpreted as having meanings consistent with their meanings in the context of the relevant art, and not as idealized or highly formalized, unless expressly defined herein.

[0038] Techniques, methods, and equipment known to those skilled in the art may not be discussed in detail, but where appropriate, they should be considered part of the specification.

[0039] like Figure 1 As shown, when the engine body 10 is running, it draws in air from the front, heats and accelerates the air, and then exhausts it to the rear. The test chamber space is divided into an inner intake area A, an outer intake area B, an engine accessory area C, an inner exhaust area E, and an outer exhaust area D. A schematic diagram of the division method is shown below. Figure 1 The aircraft engine fault monitoring system classifies the degree of fault danger based on the location of the foreign object and takes corresponding measures according to the degree of danger.

[0040] The engine body 10 contains high-speed rotating components such as a compressor and turbine. During operation, the engine body 10 may ingest intake foreign objects 6. These rotating components may fail and shatter due to high temperature, excessive speed, fatigue, or other reasons, and the fragments are carried out of the engine by the high-speed airflow, becoming exhaust splash foreign objects 8. The engine body 10 is equipped with engine accessories 5; external foreign objects that come into contact with these accessories become accessory splash foreign objects 7.

[0041] Some embodiments of this disclosure provide an aircraft engine fault monitoring system, including: an engine body 10 and one or more laser positioning radars 1, the laser positioning radars 1 being disposed around the engine body 10 and configured to detect the size parameters, position parameters and motion parameters of foreign objects.

[0042] In this illustrative embodiment, by arranging a laser positioning radar 1 around the engine body 10, the laser positioning radar 1 can accurately detect the size, position, and motion parameters of foreign objects, significantly reducing the negligence and misjudgment caused by manual monitoring of parameters, and improving the accuracy of engine fault detection. In some embodiments, such as Figure 1 As shown, the foreign objects include intake foreign objects 6, accessory splash foreign objects 7, and exhaust splash foreign objects 8.

[0043] To ensure the effectiveness of the acquired signals, in some embodiments, multiple laser positioning radars 1 are distributed at different locations around the engine body 10 to achieve redundancy detection. Redundancy detection means that the detection range of multiple laser positioning radars 1 can cover the engine body 10, and there is overlap between the detection ranges of each laser positioning radar 1.

[0044] In some embodiments, the laser positioning radar 1 includes a first laser probe and a second laser probe. The first laser probe is configured to detect and record the angular coordinates α1 and β1 of the foreign object relative to itself, and the second laser probe is configured to detect and record the angular coordinates α2 and β2 of the foreign object relative to itself. Before use, the first laser probe is calibrated to have three-dimensional absolute coordinates X1, Y1, and Z1 in space. When a foreign object is detected, the first laser probe records the angular coordinates α1 and β1 of the foreign object relative to itself. The corresponding second laser probe has its own calibrated coordinates X2, Y2, and Z2, and records the foreign object information α2 and β2. Using the above ten parameters as independent variables, the three-dimensional absolute coordinates of the foreign object can be calculated: (X 异物 Y 异物 Z 异物 ) = f(X1, Y1, Z1, α1, β1, X2, Y2, Z2, α2, β2), further improving the accuracy of foreign object detection.

[0045] like Figure 1As shown, in some embodiments, the aircraft engine fault monitoring system also includes one or more acoustic wave collectors 2 disposed around the engine body 10, configured to collect acoustic wave signals from the engine body 10. The acoustic wave collectors 2 collect local acoustic wave signals and then identify abnormal sounds during the test using a specific algorithm.

[0046] Similarly, to ensure the effectiveness of the acquired signals, in some embodiments, multiple acoustic wave acquisition devices 2 are distributed in different locations around the engine body 10 to achieve redundancy detection.

[0047] like Figure 1 As shown, in some embodiments, at least one acoustic wave collector 2 is positioned at a 60° angle to the engine axis of the engine body 10. Practice has proven that positioning the acoustic wave collector 2 at a 60° angle to the engine axis of the engine body 10 results in the strongest engine sound wave intensity at that angle, and the sound wave intensity changes significantly during a fault. Therefore, the sound intensity collected by the acoustic wave collector 2 is the most reliable and accurate, demonstrating high feasibility.

[0048] In some embodiments, the aircraft engine fault monitoring system further includes a digital controller 3 and a host computer 4. The host computer 4 is configured to receive signals from the laser positioning radar 1 and the acoustic wave collector 2 and perform data processing. The digital controller 3 is configured to control the fuel supply and adjustable actuation mechanism of the engine body 10 according to the data processed by the host computer 4, so as to realize engine speed control and shutdown operation.

[0049] The signals collected by the laser positioning radar 1 and the acoustic wave collector 2 are transmitted to the host computer 4 via wireless signals. The host computer 4 processes the signals and monitors the inhalation of foreign objects and structural failures in real time. If a fault is detected, the host computer 4 sends a fault code to the digital controller 3. The host computer 4 has two working modes: (1) Automatic mode: When the engine starts running, the engine digital controller 3 sends a start signal to the host computer 4, and the monitoring system starts automatically; (2) Manual mode: The operator manually operates the software on the host computer 4 to start the monitoring system.

[0050] When the aircraft engine fault monitoring system is activated, the laser positioning radar 1 monitors foreign objects in the entire engine test area in real time. When a foreign object is detected, the host computer 4 issues an alarm (which may include display and sound alarms) and displays the size, trajectory and speed of the foreign object on the screen of the host computer 4.

[0051] Here is a set of typical fault handling modes:

[0052] The laser positioning radar 1 detects an air intake foreign object 6 in the air intake area A. If a foreign object is found in this area, it is highly likely to be sucked into the air intake. If the diameter of the air intake foreign object 6 is >0.5cm or the speed is >50m / s, the machine will stop. Otherwise, the machine will be slowed down and stopped, and the on-site monitoring personnel will determine the subsequent operation.

[0053] The laser positioning radar 1 detects an air intake foreign object 6 in the air intake area B of the outer duct. If a foreign object is found in this area, it is highly likely to be sucked into the outer duct. If the diameter of the air intake foreign object 6 is >2cm or the speed is >100m / s, the engine will stop. Otherwise, the engine will be slowed down and stopped, and the on-site monitoring personnel will determine the subsequent operation.

[0054] Laser positioning radar 1 detects foreign object 7 splashed in the engine accessory area C. Foreign objects found in this area are likely to be foreign objects that have caused accessory failure. If the diameter of foreign object 7 splashed in the engine is >0.5cm or the speed is >50m / s, the engine should be stopped. Otherwise, the engine should be slowed down and stopped, and the on-site monitoring personnel should determine the subsequent operation.

[0055] Laser positioning radar 1 detects exhaust splash foreign object 8 in the internal exhaust zone E. The foreign object found in this area is most likely a foreign object splashed out due to the failure of the internal component. If the trajectory passes through the engine body 10, it is dealt with; otherwise, only a warning is issued. If the diameter of the exhaust splash foreign object 8 is >0.2cm and the speed is >100m / s, the engine is stopped. In other cases, the engine is slowed down and stopped, and the on-site monitoring personnel determine the subsequent operation.

[0056] Laser positioning radar 1 detects exhaust splash foreign object 8 in the exhaust area D of the outer bypass. The foreign object found in this area is most likely a foreign object splashed out due to the failure of the outer bypass component. If the trajectory passes through the engine body 10, it is dealt with; otherwise, only a warning is issued. If the diameter of the exhaust splash foreign object 8 is >0.5cm and the speed is >100m / s, the engine is stopped. In other cases, the engine is slowed down and stopped, and the on-site monitoring personnel determine the subsequent operation.

[0057] The frequency and intensity of the sound waves emitted by an engine during normal operation exhibit clear regularity, such as... Figure 2 The sound wave intensity at different frequencies was measured by sound wave collector 2 at a position 15m away from the engine axis at a 60° angle. The spike signal is N times the fan speed multiplied by the number of blades, also known as the fan blade passing frequency. That is, when the blades pass through a fixed circumferential position, they will excite the sound wave intensity of sound wave collector 2 to increase. Convert to the frequency domain diagram ( Figure 2 This means that the sound wave intensity at frequencies that are integer multiples of the fan blades passing through is represented by a spike signal. When the fan blade passing frequency and the fan rotation frequency (the reciprocal of the fan speed) overlap, the signal spike will be higher than the spike at the passing frequency of a normal blade.

[0058] Figure 3The diagram illustrates the frequency domain intensity characteristics of the fan during normal operation in some embodiments. When the compression components (fan, compressor) stall or the rotating blades rub against each other, the sound wave intensity in a certain frequency range will be abnormally high, significantly exceeding the normal operating sound wave intensity characteristics. Therefore, the aforementioned faults can be identified and avoided by analyzing the sound wave intensity. Considering the regularity of the frequency domain intensity characteristics, the intensity limits near the fan blade passing frequency and the fan rotation frequency should be given separately. The following is a table of sound wave intensity limits for a certain engine, where ±10Hz is an adjustable parameter.

[0059] Table 1. Sound Wave Intensity Limits (dB)

[0060] relative speed % Fan speed within ±10Hz Fan operates within ±10Hz Other frequencies 60％ 95 70 60 65％ 100 75 65 70％ 95 68 57 75％ 95 68 57 80％ 100 73 50 85％ 130 80 75 93％ 150 110 100 97％ 150 110 100 100％ 150 110 100 105％ 150 110 100

[0061] Table 1 only provides single-point limit values ​​for certain given rotational speeds. Between given rotational speeds, the limit values ​​are interpolated using adjacent linear interpolation (other interpolation methods, such as two-dimensional or spline interpolation, can also be used). After linear interpolation, Table 1 can be intuitively converted into... Figure 4 .

[0062] The sound wave collector 2 continuously collects audio signals from the engine's operation. When the intensity of a certain frequency exceeds... Figure 4 When the engine reaches the limit value (the adjustable limit parameter for reducing RPM to idle speed), an alarm signal is immediately sent to the host computer 4. The host computer 4 will then control the engine to reduce its RPM to idle speed via the engine's digital controller 3. If the sound wave intensity exceeds the limit during the RPM reduction process... Figure 2 The engine is stopped when the normal operating value is 1.5 times (the parking limit parameter is adjustable).

[0063] The total sound wave intensity is obtained by integrating and averaging the intensities of each frequency. Figure 5 This diagram illustrates the change in total sound wave intensity of a normally operating engine as a function of engine speed. When component failure or surge occurs in the compression components (fan / compressor), the sound wave intensity increases at all frequencies, resulting in a sharp increase in the total sound wave intensity. Sound wave acquisition device 2 continuously collects audio signals from the engine's operation; when the total sound wave intensity exceeds... Figure 5 When the engine speed reaches 1.3 times the normal operating intensity (the adjustable limit parameter for reducing RPM to idle), an alarm signal is immediately sent to the host computer 4. The host computer 4 will then control the engine to reduce its RPM to idle via the engine digital controller 3. If the total sound wave intensity exceeds [a certain value] during the RPM reduction process, [further action will be taken]. Figure 5 The engine is stopped when the normal operating value is 1.5 times (the parking limit parameter is adjustable).

[0064] The above describes the fault diagnosis logic of acoustic wave collector 2 positioned at a 60° angle to the engine axis. The engine acoustic wave intensity is strongest at this 60° angle, and the intensity changes significantly when a fault occurs. Therefore, it is recommended to place at least one acoustic wave collector 2 at this 60° angle. Differences in acoustic wave intensity collected by acoustic wave collectors 2 at other angles are described in [link to documentation]. Figure 6 It can be seen that the sound wave intensity of the engine is different at different angles, so the limit value of the sound wave collector 2 should also be different at different angles.

[0065] If the acoustic wave collector 2 is not positioned at a 60° angle relative to the engine axis, the limit values ​​can be corrected according to Table 2 below. Multiplying each limit parameter in Table 1 by the coefficient in Table 2 will yield the corresponding limit values ​​for that angle.

[0066] Table 2 Correction Table for Sound Wave Intensity Limitation Angle

[0067] Arrangement angle Limit value correction factor 120° 0.72 90° 0.91 60° 1 30° 0.85 0° 0.79

[0068] Based on the above fault detection and control logic, some embodiments of this disclosure provide a control method for controlling the aforementioned aero-engine fault monitoring system, including:

[0069] When the laser positioning radar 1 detects a foreign object in the inner air intake zone A, if the diameter of the foreign object is greater than the first preset diameter or its speed is greater than the first preset speed, it controls the engine body 10 to stop; otherwise, it controls the engine body 10 to reduce its speed to slow.

[0070] When the laser positioning radar 1 detects a foreign object in the outer bypass air intake area B, if the diameter of the foreign object is greater than the second preset diameter or its speed is greater than the second preset speed, it controls the engine body 10 to stop; otherwise, it controls the engine body 10 to reduce its speed to slow.

[0071] When the laser positioning radar 1 detects a foreign object in the accessory area C, if the diameter of the foreign object is greater than the third preset diameter or its speed is greater than the third preset speed, the engine body 10 is controlled to stop; otherwise, the engine body 10 is controlled to reduce its speed to slow.

[0072] When the laser positioning radar 1 detects a foreign object in the exhaust zone E, if the diameter of the foreign object is greater than a fourth preset diameter and its speed is greater than a fourth preset speed, the engine body 10 is stopped; otherwise, the engine body 10 is slowed down to idle speed; and / or

[0073] When the laser positioning radar 1 detects a foreign object in the exhaust area D of the outer casing, if the diameter of the foreign object is greater than the fifth preset diameter and its speed is greater than the fifth preset speed, it controls the engine body 10 to stop; otherwise, it controls the engine body 10 to reduce its speed to slow.

[0074] In some embodiments, the first preset diameter is configured to be 0.5 cm; the first preset speed is configured to be 50 m / s; in some embodiments, the second preset diameter is configured to be 2 cm; the second preset speed is configured to be 100 m / s; in some embodiments, the third preset diameter is configured to be 0.5 cm; the third preset speed is configured to be 50 m / s; in some embodiments, the fourth preset diameter is configured to be 0.2 cm; the fourth preset speed is configured to be 100 m / s; in some embodiments, the fifth preset diameter is configured to be 0.5 cm; the fifth preset speed is configured to be 100 m / s.

[0075] Similarly, based on the above-described fault detection and control logic, some embodiments of this disclosure provide a control method for controlling the aforementioned aero-engine fault monitoring system, including:

[0076] If the intensity of the sound wave collected by the sound wave collector 2 is between a first preset intensity value and a second preset intensity value, the engine body 10 is controlled to stop; and

[0077] If the intensity of the sound wave collected by the sound wave collector 2 exceeds the second preset intensity value, the engine body 10 will be stopped.

[0078] The first preset strength value and the second preset strength value can be set according to different parameters of the engine. The relevant embodiments have been given in the above control logic, and will not be repeated here.

[0079] The embodiments of this disclosure have now been described in detail. To avoid obscuring the concept of this disclosure, some details known in the art have not been described. Those skilled in the art can fully understand how to implement the technical solutions disclosed herein based on the above description.

[0080] While specific embodiments of this disclosure have been described in detail by way of examples, those skilled in the art should understand that the examples are for illustrative purposes only and not intended to limit the scope of this disclosure. Those skilled in the art should understand that modifications can be made to the above embodiments or equivalent substitutions can be made to some technical features without departing from the scope and spirit of this disclosure. The scope of this disclosure is defined by the appended claims.

Claims

1. A control method applied to an aircraft engine fault monitoring system, characterized in that, The aircraft engine fault monitoring system includes: Engine body (10); and One or more laser positioning radars (1) are disposed around the engine body (10) and are configured to detect the size parameters, position parameters and motion parameters of foreign objects; The control method includes: When the laser positioning radar (1) detects a foreign object in the inner air intake area (A), if the diameter of the foreign object is greater than the first preset diameter or its speed is greater than the first preset speed, it controls the engine body (10) to stop; otherwise, it controls the engine body (10) to reduce its speed to slow. When the laser positioning radar (1) detects a foreign object in the outer bypass air intake area (B), if the diameter of the foreign object is greater than the second preset diameter or its speed is greater than the second preset speed, it controls the engine body (10) to stop; otherwise, it controls the engine body (10) to reduce its speed to slow. When the laser positioning radar (1) detects a foreign object in the accessory area (C), if the diameter of the foreign object is greater than the third preset diameter or its speed is greater than the third preset speed, it controls the engine body (10) to stop; otherwise, it controls the engine body (10) to reduce its speed to slow. When the laser positioning radar (1) detects a foreign object in the inner exhaust zone (E), if the diameter of the foreign object is greater than a fourth preset diameter and its speed is greater than a fourth preset speed, it controls the engine body (10) to stop; otherwise, it controls the engine body (10) to reduce its speed to idle. When the laser positioning radar (1) detects a foreign object in the outer exhaust area (D), if the diameter of the foreign object is greater than the fifth preset diameter and its speed is greater than the fifth preset speed, it controls the engine body (10) to stop; otherwise, it controls the engine body (10) to reduce its speed to slow.

2. The control method according to claim 1, characterized in that, The foreign objects include air intake foreign objects (6), accessory splash foreign objects (7), and exhaust splash foreign objects (8).

3. The control method according to claim 1, characterized in that, Multiple laser positioning radars (1) are distributed in different locations around the engine body (10) to achieve redundant detection.

4. The control method according to claim 1, characterized in that, The laser positioning radar (1) includes a first laser probe and a second laser probe. The first laser probe is configured to detect and record the angular coordinates α1 and β1 of the foreign object relative to the first laser probe, and the second laser probe is configured to detect and record the angular coordinates α2 and β2 of the foreign object relative to the second laser probe.

5. The control method according to any one of claims 1 to 4, characterized in that, The aircraft engine fault monitoring system also includes one or more acoustic wave collectors (2) disposed around the engine body (10) and configured to collect acoustic wave signals from the engine body (10).

6. The control method according to claim 5, characterized in that, Multiple acoustic wave collectors (2) are distributed in different locations around the engine body (10) to achieve redundancy detection.

7. The control method according to claim 5, characterized in that, At least one of the acoustic wave collectors (2) is positioned at a 60° angle to the engine axis of the engine body (10).

8. The control method according to claim 5, characterized in that, The aircraft engine fault monitoring system also includes a digital controller (3) and a host computer (4). The host computer (4) is configured to receive signals from the laser positioning radar (1) and the acoustic wave collector (2) and perform data processing. The digital controller (3) is configured to control the fuel supply and adjustable actuation mechanism of the engine body (10) according to the data processed by the host computer (4) to realize engine speed control and shutdown operation.

9. The control method according to claim 8, characterized in that, include: If the intensity of the sound wave collected by the sound wave collector (2) is between the first preset intensity value and the second preset intensity value, the engine body (10) is controlled to stop. and If the intensity of the sound wave collected by the sound wave collector (2) exceeds the second preset intensity value, the engine body (10) is controlled to stop.

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