Aero-engine state monitoring method and device and computer readable storage medium

Abstract

The invention discloses an aero-engine state monitoring method and device and a computer readable storage medium, and relates to the technical field of aero-engine state monitoring. The method comprises the following steps: acquiring a magnetic flux variable quantity, an optical signal, a thermal radiation signal and a charge variable quantity through a composite sensor, and converting the magnetic flux variable quantity, the optical signal, the thermal radiation signal and the charge variable quantity into coil output voltage and optical signal voltage, and converting into thermal radiation exitance voltage signals and variable voltage under different wavelengths; obtaining a blade tip gap measured value according to the coil output voltage and the distance value between the probe and the blade tip; based on the blade tip timing principle, blade vibration displacement is obtained according to the optical signal voltage; performing colorimetric processing on the thermal radiation exitance voltage signals under different wavelengths to obtain a radiation exitance ratio at different wavelengths, and further obtaining the temperature of a turbine blade temperature field; and calculating the casing acceleration according to the changing voltage. In conclusion, state parameters in multiple aspects are obtained at the same measuring point through one composite sensor, and the engineering testing steps are simplified.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aero-engine state monitoring, and particularly relates to an aero-engine state monitoring method, device and computer readable storage medium. BACKGROUND

[0002] As the core equipment of an aircraft, an aero-engine plays an irreplaceable role in air transportation and national defense construction. It is crucial to deeply explore the fault prediction and health management technology of the aero-engine. The blade is the most numerous and weakest core component in the aero-engine. In order to improve the fuel efficiency of the aero-engine, ensure the operation safety and prolong the service life of the engine, it is necessary to accurately measure the blade tip clearance, blade tip vibration, blade surface temperature and casing acceleration.

[0003] However, in the existing measurement device and method, one probe can only measure one or two physical quantities. Since there are a large number of components installed on the casing, the positions where the sensors can be arranged are limited, and it is difficult to realize the multi-aspect state monitoring of the aero-engine. SUMMARY

[0004] Therefore, the present application aims to overcome the deficiencies in the prior art, and provides an aero-engine state monitoring method, device and computer readable storage medium, which are used for monitoring the working state parameters of the rotating blades of the aero-engine and providing more reliable information for the health management and fault prediction of the engine.

[0005] The present application provides the following technical solutions: In a first aspect, the present application provides an aero-engine state monitoring method, comprising: obtaining a magnetic flux change amount, an optical signal, a thermal radiation signal and a charge change amount through a composite sensor; converting the magnetic flux change amount into a coil output voltage, converting the optical signal into an optical signal voltage, converting the thermal radiation signal to obtain a thermal radiation exitance voltage signal at different wavelengths, and converting the charge change amount into a change voltage; obtaining a measured value of the blade tip clearance according to the coil output voltage and the distance between the probe and the blade tip; obtaining a blade vibration displacement according to the optical signal voltage based on the blade tip timing principle; performing colorimetric processing on the thermal radiation exitance voltage signal at different wavelengths to obtain a radiation exitance ratio at different wavelengths, and obtaining a turbine blade temperature field temperature according to the radiation exitance ratio at different wavelengths; calculating the casing acceleration according to the change voltage.

[0006] In an embodiment, the step of obtaining a measured value of the blade tip clearance according to the coil output voltage and the distance between the probe and the blade tip comprises: obtaining a preset mapping relationship between the coil output voltage experimental value and the probe-to-tip clearance experimental value; based on the preset mapping relationship, calculating the tip clearance measured value according to the coil output voltage.

[0007] In an embodiment, the blade vibration displacement is obtained according to the light signal voltage based on the tip timing principle, comprising: when the voltage amplitude of the light signal voltage reaches a preset threshold value, the time corresponding to the current light signal voltage is determined as the blade arrival time stamp; calculating the blade tip line speed according to the preset key phase time stamp; calculating the blade vibration displacement according to the blade tip line speed, the preset key phase time stamp, the blade arrival time stamp, and the preset arc length between the blade and the key phase.

[0008] In an embodiment, the turbine blade temperature field temperature is obtained according to the radiation emittance ratio at different wavelengths, comprising: obtaining a numerical relationship between the radiation emittance ratio at different wavelengths and the object temperature; based on the classical Planck blackbody radiation law, determining the radiation emittance ratio formula of the object at a preset temperature at wavelengths 、 , , wherein is the preset temperature, are the radiation emittance of the object at wavelengths 、 , is the second radiation constant; and 、 are close, , the real temperature of the object is: , wherein only is a variable; obtaining the turbine blade temperature field temperature according to the radiation emittance ratio at different wavelengths and the radiation emittance ratio formula, and the calculation formula of the turbine blade temperature field temperature is: .

[0009] In an embodiment, the machine case acceleration is calculated according to the change voltage, comprising: obtaining a preset numerical relationship between the charge sensitivity experimental value, the feedback capacitance experimental value of the charge quantity conversion stage, the change voltage experimental value, and the machine case acceleration experimental value; the expression of the preset numerical relationship is: , is the change voltage experimental value, an experimental value of the feedback capacitance of the charge quantity conversion stage, an experimental value of the charge sensitivity, a an experimental value of the casing acceleration, calculating the casing acceleration according to the change voltage based on the preset numerical relationship.

[0010] In a second aspect, the present application provides an aero-engine state monitoring device for executing the aero-engine state monitoring method of the first aspect, the device comprising: a composite sensor, an eddy current preamplifier, a vibration photoelectric conversion module, a temperature photoelectric conversion module, a charge amplifier, and a signal acquisition and data processing system. The composite sensor is configured to acquire a magnetic flux change, an optical signal, a thermal radiation signal, and a charge change. The eddy current preamplifier is configured to convert the magnetic flux change into a coil output voltage and output the coil output voltage to the signal acquisition and data processing system. The vibration photoelectric conversion module is configured to convert the optical signal into an optical signal voltage and output the optical signal voltage to the signal acquisition and data processing system. The temperature photoelectric conversion module is configured to convert the thermal radiation signal at different wavelengths and output the thermal radiation signal at different wavelengths to the signal acquisition and data processing system. The charge amplifier is configured to convert the charge change into a change voltage and output the change voltage to the signal acquisition and data processing system. The signal acquisition and data processing system is configured to obtain a measured value of the tip clearance based on the coil output voltage and the value of the distance between the probe and the tip; obtain the vibration displacement of the blade based on the optical signal voltage based on the tip timing principle; perform colorimetric processing on the thermal radiation exitance voltage signal at different wavelengths to obtain the radiation exitance ratio at different wavelengths, and further obtain the temperature field temperature of the turbine blade; and calculate the casing acceleration based on the change voltage.

[0011] In an embodiment, the composite sensor comprises a sensor probe, a signal splitter, an emission optical fiber, an eddy current signal line, a temperature receiving optical fiber, a vibration receiving optical fiber, and an acceleration signal line. The sensor probe is installed on a composite sensor casing. The signal splitter is configured to transmit the magnetic flux change to the eddy current signal line, transmit the optical signal to the vibration receiving optical fiber, transmit the thermal radiation signal to the temperature receiving optical fiber, and transmit the charge change to the acceleration signal line.

[0012] In an embodiment, the sensor probe comprises a signal line transmission channel connecting the transmitting optical fiber, the temperature receiving optical fiber, the vibration receiving optical fiber, the eddy current signal line and the acceleration signal line.

[0013] In an embodiment, the sensor probe further comprises a composite sensor shell and a piezoelectric module, the piezoelectric module being a nut-type piezoelectric module mounted on the composite sensor shell through threaded connection.

[0014] In a third aspect, the present application provides a computer readable storage medium storing a computer program, which, when executed by a processor, implements the aero-engine state monitoring method according to the first aspect.

[0015] The aero-engine state monitoring method, device and computer readable storage medium disclosed by the present application obtain the magnetic flux variation, the light signal, the thermal radiation signal and the charge variation through a composite sensor; convert the magnetic flux variation into a coil output voltage, convert the light signal into a light signal voltage, convert the thermal radiation signal to obtain a thermal radiation emittance voltage signal at different wavelengths, and convert the charge variation into a change voltage; obtain a blade tip clearance measured value according to the coil output voltage and the distance between the probe and the blade tip; obtain a blade vibration displacement according to the light signal voltage based on the blade tip timing principle; perform colorimetric processing on the thermal radiation emittance voltage signal at different wavelengths to obtain a radiation emittance ratio at different wavelengths, and further obtain a turbine blade temperature field temperature; and calculate the casing acceleration according to the change voltage. In summary, the composite sensor simultaneously obtains the blade tip clearance measured value, the blade vibration displacement, the turbine blade temperature field temperature and the casing acceleration and other state parameters at the same measurement point, simplifies the engineering test steps, and provides a reliable monitoring method for aero-engine health management and fault prediction. BRIEF DESCRIPTION OF DRAWINGS

[0016] In order to more clearly illustrate the technical solutions of the present application, the following will briefly introduce the drawings needed to be used in the embodiments. It should be understood that the following drawings only show some embodiments of the present application, and therefore should not be regarded as limiting the scope of protection of the present application. In each drawing, similar components are denoted by similar reference numerals.

[0017] Figure 1 A flowchart of the aero-engine state monitoring method proposed in the present embodiment is shown; Figure 2 A structural diagram of the aero-engine state monitoring device proposed in the present embodiment is shown; Figure 3 A structural diagram of the composite sensor proposed in the present embodiment is shown; Figure 4 A structural schematic diagram of a sensor probe proposed in the embodiment is shown; Figure 5 A schematic diagram of internal layout of a fiber ferrule proposed in the embodiment is shown; Figure 6 A schematic diagram of a piezoelectric module mounting method proposed in the embodiment is shown.

[0018] Explanation of the drawing: 201-composite sensor; 202-laser emitter; 203-electric eddy current preamplifier; 204-temperature photoelectric conversion module; 205-vibration photoelectric conversion module; 206-charge amplifier; 207-signal acquisition and data processing system; 2011-sensor probe; 2012-armor protection shell; 2013-signal splitter; 2014-emitting optical fiber; 2015-eddy current signal line; 2016-temperature receiving optical fiber; 2017-vibration receiving optical fiber; 2018-acceleration signal line; 401-eddy current probe shell; 402-eddy current coil; 403-coil shell; 404-coil support; 405-fiber ferrule; 406-piezoelectric module; 407-signal line transmission channel; 408-mounting hole; 409-composite sensor shell; 601-nut type piezoelectric module. DETAILED DESCRIPTION

[0019] The technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments of the present application.

[0020] The components of the embodiments of the present application generally described and shown in the drawings herein can be arranged and designed in various different configurations. Therefore, the following detailed description of the embodiments of the present application provided in the drawings is not intended to limit the scope of the claimed application, but only represents selected embodiments of the present application. Based on the embodiments of the present application, all other embodiments obtained by those skilled in the art without creative work are within the scope of protection of the present application.

[0021] Hereinafter, the terms "include", "have", and their conjugates, used in various embodiments of the present application, are only intended to denote that specific features, numbers, steps, operations, elements, components, or combinations thereof are present, and should not be understood as excluding the presence or addition of one or more other features, numbers, steps, operations, elements, components, or combinations thereof.

[0022] In addition, the terms "first", "second", "third", and the like are only used to distinguish descriptions, and cannot be understood as indicating or implying relative importance.

[0023] Unless otherwise defined, all terms (including technical and scientific terms) used herein have the same meaning as commonly understood by one of ordinary skill in the art to which various embodiments of the present application belong. It will be further understood that terms, such as those defined in commonly used dictionaries, should be interpreted as having a meaning that is consistent with their meaning in the context of the relevant art and will not be interpreted in an overly idealized or formal sense unless expressly so defined in various embodiments of the present application.

[0024] Embodiment 1 The embodiment of the present disclosure provides an aero-engine state monitoring method, which is applied to an aero-engine state monitoring device and used for monitoring a rotating aero-engine blade working state parameter, so as to provide more reliable information for engine health management and fault prediction.

[0025] Please refer to Figure 1 , the aero-engine state monitoring method comprises steps S101-S106, and each step will be described in detail below.

[0026] In step S101, the magnetic flux variation, the light signal, the thermal radiation signal and the charge variation are obtained by the composite sensor.

[0027] In the embodiment, please refer to Figure 2 , the aero-engine state monitoring device comprises a composite sensor 201 and a laser emitter 202.

[0028] When monitoring the rotating aero-engine blade working parameter, the laser emitter 202 generates an excitation magnetic field in the direction of the blade and emits a light signal of a specific wavelength, and the composite sensor 201 obtains the magnetic flux variation, the light signal, the thermal radiation signal and the charge variation when the blade passes.

[0029] Please refer to Figure 3 , the composite sensor 201 comprises a sensor probe 2011, an armored protective shell 2012, a signal splitter 2013, a transmitting optical fiber 2014, an eddy current signal line 2015, a temperature receiving optical fiber 2016, a vibration receiving optical fiber 2017 and an acceleration signal line 2018. The sensor probe 2011 is installed on the casing; the signal splitter 2013 transmits the measured magnetic flux variation, light signal, thermal radiation signal and charge variation into four groups of corresponding receivers, the transmitting optical fiber 2014 is used to transmit the light signal generated by the laser emitter 202 to the blade surface, the temperature receiving optical fiber 2016 and the vibration receiving optical fiber 2017 respectively receive the light signal and the thermal radiation signal of the scattered light returned by the blade surface, and the acceleration signal line 2018 is used to receive the charge variation generated by the piezoelectric sensor assembly.

[0030] It should be noted that the sensor probe 2011 can be installed at any position of the blade circumference, and the blade will trigger the sensor to generate a corresponding induction signal when passing through the blade. The tail joint of the transmitting optical fiber 2014, the temperature receiving optical fiber 2016 and the vibration receiving optical fiber 2017 can adopt LC, SC, FC, ST and other forms, and the eddy current signal line 2015 and the acceleration signal line 2018 can adopt SMA, QMA, BMA and other forms.

[0031] It should be noted that please refer to Figure 4 The sensor probe 2011 includes an eddy current probe shell 401, an eddy current coil 402, a coil shell 403, a coil support 404, an optical fiber ferrule 405, a piezoelectric module 406, a signal line transmission channel 407, a mounting hole 408 and a composite sensor shell 409.

[0032] Specifically, the material of the eddy current probe shell 401 is not limited and can be stainless steel. The optical fiber ferrule 405 is a hollow sleeve with a guide groove, which is glued with the coil support 404 using epoxy resin. The eddy current coil 402 is wound on the coil support 404. The eddy current probe shell 401 and the coil shell 403 are glued with other parts using epoxy resin. The piezoelectric module 406 is connected with the eddy current probe shell 401 at the mounting hole 408 by screw thread. The piezoelectric module 406 is glued with the signal line transmission channel 407 using epoxy resin. Five channels are provided in the middle of the signal line transmission channel 407, which are connected with the transmitting optical fiber 2014, the temperature receiving optical fiber 2016, the vibration receiving optical fiber 2017, the eddy current signal line 2015 and the acceleration signal line 2018 respectively.

[0033] Specifically, the mounting hole 408 is fixed with the piezoelectric module 406 by a double-nut anti-loose structure. The nut can be a nylon insert nut, a threaded locking nut or a spring washer to increase the anti-vibration and anti-loose ability. This structure is reliable and easy to disassemble.

[0034] The composite sensor shell 409 is a hollow stainless steel shell with external threads. A layer of silicon oxide protective film is evaporated on the top end of the outer wall to prevent metal evaporation in high temperature environment. The hollow stainless steel shell of the sensor probe 2011 can be provided with an air cooling sleeve, which is matched with a cooling gas source to realize real-time cooling of the sensor by introducing external cooling gas.

[0035] Please refer to Figure 5, the optical fiber ferrule 405 adopts a full optical fiber structure, the transmitting optical fiber 2014, the temperature receiving optical fiber 2016 and the vibration receiving optical fiber 2017 are arranged in the optical fiber ferrule 405. The transmitting optical fiber 2014 is one optical fiber, the temperature receiving optical fiber 2016 and the vibration receiving optical fiber 2017 are two or more optical fibers. For example, the transmitting optical fiber 2014 is one optical fiber, the temperature receiving optical fiber 2016 is two optical fibers, and the vibration receiving optical fiber 2017 is four optical fibers; in the optical fiber ferrule 405, the transmitting optical fiber 2014 is located at the center of the optical fiber bundle, the temperature receiving optical fiber 2016 and the vibration receiving optical fiber 2017 are distributed around the transmitting optical fiber 2014 in the form of concentric circles, and the two temperature receiving optical fibers 2016 are symmetrically arranged.

[0036] Please refer to Figure 6 , the piezoelectric module 406 can adopt a nut type piezoelectric module 601, which is installed on the composite sensor shell 409 through threaded connection, in order to prevent the nut type piezoelectric module 601 from loosening, double nut anti-loosening structure, installation spring washer and thread paste can be selected to prevent loosening, the embodiment has the advantages of convenient disassembly and convenient replacement of internal components of the probe.

[0037] Step S102, converting the magnetic flux change into a coil output voltage, converting the optical signal into an optical signal voltage, converting the thermal radiation signal to obtain a thermal radiation exitance voltage signal at different wavelengths, and converting the charge change into a change voltage.

[0038] In this embodiment, please refer again to Figure 2 , the aero-engine state monitoring device further comprises an eddy current preamplifier 203, a temperature photoelectric conversion module 204, a vibration photoelectric conversion module 205, a charge amplifier 206, and a signal acquisition and data processing system 207.

[0039] The eddy current preamplifier 203 receives the magnetic flux change in the eddy current coil and converts the magnetic flux change into a coil output voltage, and then outputs the coil output voltage to the signal acquisition and data processing system 207.

[0040] The temperature photoelectric conversion module 204 is connected with the temperature receiving optical fiber 2016, the temperature receiving optical fiber 2016 transmits the received thermal radiation signal to the temperature photoelectric conversion module 204 for processing, converts the thermal radiation signal into a thermal radiation exitance voltage signal at different wavelengths, and then sends the thermal radiation signal voltage into the signal acquisition and data processing system 207.

[0041] The vibration photoelectric conversion module 205 is connected with the vibration receiving optical fiber 2017, the vibration receiving optical fiber 2017 transmits the received optical signal to the vibration photoelectric conversion module 205 for processing, converts the optical signal into an optical signal voltage, and then sends the optical signal voltage into the signal acquisition and data processing system 207.

[0042] The acceleration signal line 2018 is connected with the charge amplifier 206, the acceleration signal line 2018 converts the received charge variation of the piezoelectric sensor assembly into a variation voltage, and then sends the variation voltage into the signal acquisition and data processing system 207.

[0043] It should be noted that the eddy current preamplifier 203 comprises a high-frequency signal generator, a signal conditioning circuit and a linearization processing module. The high-frequency signal generator generates a high-frequency alternating current signal to drive the probe coil to generate an alternating magnetic field to excite the eddy current effect on the surface of the measured conductor. The demodulation circuit in the signal conditioning circuit converts the coil impedance variation into a voltage signal, the amplification circuit amplifies the weak detection signal to improve the signal-to-noise ratio, the filter circuit filters out high-frequency noise and interference and retains the effective signal. The linearization processing module compensates for the non-linear characteristics of the eddy current sensor and outputs a linear signal proportional to the displacement. The eddy current signal is transmitted to the eddy current preamplifier 203 through the eddy current signal line 2015, converted into a voltage signal, and then input into the signal acquisition and data processing system 207.

[0044] The temperature photoelectric conversion module 204 and the vibration photoelectric conversion module 205 each comprise a photodetector and a multi-stage amplification circuit. The photodetector can be selected from various photodiodes or photomultiplier tubes according to the wavelength range of the spectrum, such as a Si-PIN photodiode. The Si-PIN photodiode is connected with the temperature receiving optical fiber 2016 or the vibration receiving optical fiber 2017 in a fiber-tail structure, and the optical signal is converted into a current signal by the diode. The multi-stage amplification circuit converts the current signal into a voltage signal and inputs the voltage signal into the signal acquisition and data processing system 207 after multi-stage amplification. In addition, the temperature photoelectric conversion module 204 further comprises a double-channel light splitting system. The thermal radiation signal is collimated, split and focused by the double-channel light splitting system to obtain two groups of thermal radiation signals with different wavelengths, which are then transmitted to the photodetector.

[0045] The charge amplifier 206 comprises a charge converter, a signal conditioning circuit and an adjustable gain amplifier. The charge converter converts the high-impedance charge signal output by the piezoelectric sensor into a voltage signal; the signal conditioning circuit eliminates the direct current drift of the piezoelectric sensor and suppresses noise; the adjustable gain amplifier provides multi-stage gain adjustment to adapt to sensor signals of different ranges and ensure that the output signal matches the input range of the subsequent acquisition equipment; the acceleration signal is transmitted to the charge amplifier 206 through the acceleration signal line 2018, converted into a voltage signal, and then input into the signal acquisition and data processing system 207.

[0046] In the signal acquisition and data processing system 207, the signal acquisition part mainly includes an acquisition card and a communication module; the data processing part mainly includes a memory, a processor and a corresponding data acquisition processing program in the processor, wherein the processor executes the data processing program to calculate and analyze the multi-channel voltage signals collected by the data acquisition card to obtain the blade state information.

[0047] In step S103, the measured value of the tip clearance is obtained according to the coil output voltage and the value of the distance between the probe and the tip clearance.

[0048] In this embodiment, the eddy current sensor probe is connected to an alternating current, which forms a magnetic field around it. When the rotating blade sweeps across the sensor, it cuts the magnetic induction lines, generates an induced current, forms a magnetic field opposite to the original coil magnetic field, and changes the equivalent impedance of the coil. The expression of the equivalent impedance is: , wherein, is the electrical conductivity of the material, is the magnetic permeability of the material, is the frequency of the alternating current, is the size factor, is the tip clearance. For a known sensor and a test piece, , , , are all constant values. The coil impedance value is only related to the tip clearance, so the measured value of the tip clearance can be calculated by the signal acquisition and data processing system according to the coil output voltage and the value of the distance between the probe and the tip clearance.

[0049] In a specific embodiment, step S103 includes: obtaining a preset mapping relationship between the coil output voltage experimental value and the distance between the probe and the tip clearance experimental value; and calculating the measured value of the tip clearance according to the coil output voltage based on the preset mapping relationship.

[0050] In this embodiment, the preset mapping relationship between the coil output voltage experimental value and the distance between the probe and the tip clearance experimental value of the eddy current coil is calibrated in advance by a static calibration test bench, wherein the distance between the probe and the tip clearance is taken as the independent variable and the coil output voltage is taken as the dependent variable; the rotational angle of the eddy current sensor and the tip is kept unchanged, the tip clearance is changed, and the corresponding coil output voltage under each turbine tip clearance is obtained; based on the preset mapping relationship, the measured value of the tip clearance is determined according to the coil output voltage.

[0051] It should be noted that the static calibration is carried out in a controlled environment, and the precise correspondence between the physical quantity (gap) and the electrical signal (voltage) is directly established. This effectively avoids errors introduced by factors such as sensor nonlinearity and temperature drift, significantly improving gap measurement accuracy. At the same time, the sensor and the rotation angle of the blade tip are strictly maintained during calibration and application, eliminating the influence of the installation angle change on the vortex field distribution, ensuring the uniqueness and reliability of the mapping relationship; and during measurement, only the coil output voltage needs to be read, and the measured value of the current gap can be obtained according to the preset mapping relationship, meeting the real-time monitoring requirements of rotating machinery. In addition, the eddy current principle realizes non-wearing measurement and does not affect the dynamics of the blade, and is suitable for harsh environments such as high temperature and high speed.

[0052] In step S104, the blade vibration displacement is obtained based on the light signal voltage according to the blade tip timing principle.

[0053] In this embodiment, the signal acquisition and data processing system calculates the light signal voltage based on the blade tip timing principle to obtain the vibration displacement of the turbine blade.

[0003] In a specific embodiment, step S104 includes: when the voltage amplitude of the light signal voltage reaches a preset threshold, the time corresponding to the current light signal voltage is determined as the blade arrival timestamp; the blade tip linear velocity is calculated according to the preset key phase timestamp; and the blade vibration displacement is calculated according to the blade tip linear velocity, the preset key phase timestamp, the blade arrival timestamp, and the preset arc length between the blade and the key phase.

[0055] In this embodiment, when the voltage amplitude of the light signal voltage reaches a preset threshold, the time corresponding to the current light signal voltage is determined as the blade arrival timestamp ; the blade tip linear velocity is calculated according to the preset key phase timestamp , and the calculation formula of the blade tip linear velocity is: , wherein ; the blade vibration displacement is calculated according to the blade tip linear velocity , the preset key phase timestamp , the blade arrival timestamp , and the preset arc length between the blade and the key phase, and the calculation expression of the blade vibration displacement is: , , wherein b represents the blade number, and L represents the arc length between the blade and the key phase.

[0056] Exemplarily, when there is no key phase, assuming that blade 1 is taken as a reference blade, the blade vibration displacement n of the remaining calculation blades at the first circle of rotation can be represented as: , .

[0057] Step S105, the thermal radiation exitance signal voltage at different wavelengths is subjected to colorimetric processing to obtain the radiation exitance ratio at different wavelengths, and the turbine blade temperature field temperature is obtained according to the radiation exitance ratio at different wavelengths.

[0058] In the embodiment, the thermal radiation exitance signal voltage is proportional to the radiation exitance at different wavelengths, so the thermal radiation exitance signal voltage ratio at different wavelengths is the radiation exitance ratio at different wavelengths; further, a relationship model between the radiation exitance ratio at different wavelengths and the temperature is established to obtain the turbine blade temperature field temperature.

[0059] In a specific embodiment, step S105 comprises: obtaining a numerical relationship between the radiation exitance ratio at different wavelengths and the object temperature; determining the radiation exitance ratio formula of the object at preset temperature at wavelengths , according to the numerical relationship and the classical Planck blackbody radiation law, wherein , is the preset temperature, is the radiation emissivity of the object at wavelengths , , is the second radiation constant; and , are similar, , the real temperature of the object is: , wherein only is a variable; the turbine blade temperature field temperature is obtained according to the radiation exitance ratio at different wavelengths and the radiation exitance ratio formula, and the calculation formula of the turbine blade temperature field temperature is: .

[0060] In the embodiment, in actual turbine blade temperature measurement, the blade is not an ideal blackbody, and its radiation characteristics will be affected by factors such as material and surface state, and has different radiation emissivities at different wavelengths. The temperature measurement method considering the emissivity change can more accurately reflect the real ratio of the radiation exitance of the object at different wavelengths, and further improve the accuracy of temperature measurement. At the same time, the radiation exitance ratio formula of the object at preset temperature at wavelengths , is obtained, and the turbine blade temperature field temperature is obtained according to the radiation exitance ratio formula.The radiation exitance ratio formula accurately establishes the relationship between the radiation exitance ratio and the object temperature, reduces the error caused by approximate processing or empirical formula, and provides a reliable theoretical basis for accurately obtaining the turbine blade temperature. In addition, when calculating the temperature, only the radiation exitance ratios at two wavelengths need to be measured, and the object temperature can be quickly calculated by substituting the formula, and the calculation process is relatively simple, which is convenient for practical application and real-time monitoring. In the formula, only is a variable.

[0061] In step S106, the case acceleration is calculated according to the change voltage.

[0062] In the embodiment, the case vibration acceleration is obtained by the signal acquisition and data processing system according to the change voltage output by the piezoelectric material and the acceleration calibration curve.

[0063] In an embodiment, step S106 includes: obtaining a preset numerical relationship between a charge sensitivity experimental value, a feedback capacitance experimental value of a charge quantity conversion stage, a change voltage experimental value, and a case acceleration experimental value; an expression of the preset numerical relationship is: , is the change voltage experimental value, is the feedback capacitance experimental value of the charge quantity conversion stage, is the charge sensitivity experimental value, a is the case acceleration experimental value; the case acceleration is calculated according to the change voltage based on the preset numerical relationship.

[0064] In the embodiment, when the piezoelectric crystal is stressed, two surfaces of the crystal form a parallel plate capacitor, and the charge quantity generated on the crystal sheet is proportional to the force. The change voltage output after the charge quantity is processed by the charge amplifier is related to the following formula: , in which, is the open-loop amplification of the charge amplifier, q is the charge quantity generated on the crystal sheet, is the internal inherent capacitance of the sensor, is the distributed capacitance of the connecting cable, is the equivalent capacitance of the amplifier, is the feedback capacitance of the charge quantity conversion stage, is the change voltage. Usually , so , and the expression of the change voltage can be simplified as: , in which, is the charge sensitivity of the sensor, a is the input acceleration of the sensor, i.e., the case acceleration.

[0065] It should be noted that the physical law of direct piezoelectric sensing (charge output is proportional to acceleration) and the principle of charge amplifier circuit (output voltage is inversely proportional to input charge / feedback capacitance). This conversion based on physical model has high reliability and theoretical basis.

[0066] The aero-engine state monitoring method provided in the embodiment obtains the magnetic flux variation, the light signal, the thermal radiation signal and the charge variation through the composite sensor; converts the magnetic flux variation into a coil output voltage, converts the light signal into a light signal voltage, converts the thermal radiation signal to obtain a thermal radiation exitance voltage signal at different wavelengths, and converts the charge variation into a change voltage; obtains a blade tip clearance measured value according to the coil output voltage and the probe-to-blade tip clearance distance value; obtains a blade vibration displacement according to the light signal voltage based on the blade tip timing principle; performs colorimetric processing on the thermal radiation exitance voltage signal at different wavelengths to obtain a radiation exitance ratio at different wavelengths, and further obtains a turbine blade temperature field temperature; and calculates a casing acceleration according to the change voltage. In summary, the composite sensor is used to simultaneously obtain the blade tip clearance measured value, the blade vibration displacement, the turbine blade temperature field temperature and the casing acceleration and other state parameters at the same measurement point, thereby simplifying the engineering test steps and providing a reliable monitoring method for aero-engine health management and fault prediction.

[0067] Embodiment 2 In addition, the embodiment of the present disclosure provides an aero-engine state monitoring device for executing the aero-engine state monitoring method of the first aspect, the device comprising: a composite sensor, an eddy current preamplifier, a temperature photoelectric conversion module, a vibration photoelectric conversion module, a charge amplifier and a signal acquisition and data processing system; The composite sensor is used to obtain the magnetic flux variation, the light signal, the thermal radiation signal and the charge variation; The eddy current preamplifier is used to convert the magnetic flux variation into a coil output voltage, and output the coil output voltage to the signal acquisition and data processing system; The vibration photoelectric conversion module is used to convert the light signal into a light signal voltage, and output the light signal voltage to the signal acquisition and data processing system; The temperature photoelectric conversion module is used to convert the thermal radiation signal to obtain a thermal radiation signal at different wavelengths, and convert the thermal radiation signal at different wavelengths into a radiation exitance voltage signal and output to the signal acquisition and data processing system; The charge amplifier is used to convert the charge variation into a change voltage, and output the change voltage to the signal acquisition and data processing system; The signal acquisition and data processing system is used to obtain a measured value of the blade tip clearance according to the coil output voltage and the distance between the probe and the blade tip clearance; the blade vibration displacement is obtained according to the light signal voltage based on the blade tip timing principle; the radiation emittance ratio at different wavelengths is obtained by colorimetric processing of the thermal radiation emittance voltage signals at different wavelengths, and then the temperature field temperature of the turbine blade is obtained; and the computer case acceleration is calculated according to the change voltage.

[0068] Optionally, the composite sensor comprises a sensor probe, a signal distributor, a transmitting optical fiber, an eddy current signal line, a temperature receiving optical fiber, a vibration receiving optical fiber and an acceleration signal line; the sensor probe is installed on the composite sensor case; the signal distributor is used to transmit the magnetic flux change to the eddy current signal line, transmit the light signal to the vibration receiving optical fiber, transmit the thermal radiation signal to the temperature receiving optical fiber and transmit the charge change to the acceleration signal line.

[0069] The sensor probe comprises a signal line transmission channel, and the signal line transmission channel is connected with the transmitting optical fiber, the temperature receiving optical fiber, the vibration receiving optical fiber, the eddy current signal line and the acceleration signal line.

[0070] Optionally, the sensor probe further comprises a composite sensor shell and a piezoelectric module, and the piezoelectric module is a nut type piezoelectric module which is installed on the composite sensor shell through thread connection.

[0071] Optionally, the signal acquisition and data processing system is further used to obtain a preset mapping relationship between the coil output voltage experimental value and the distance between the probe and the blade tip clearance experimental value; and the measured value of the blade tip clearance is calculated according to the coil output voltage based on the preset mapping relationship.

[0072] Optionally, the signal acquisition and data processing system is further used to determine the time corresponding to the current light signal voltage as a blade arrival time stamp when the voltage amplitude of the light signal voltage reaches a preset threshold value; the blade tip linear velocity is calculated according to a preset key phase time stamp; and the blade vibration displacement is calculated according to the blade tip linear velocity, the preset key phase time stamp, the blade arrival time stamp and a preset arc length between the blade and the key phase.

[0073] Optionally, the signal acquisition and data processing system is further used to obtain a numerical relationship between the radiation emittance ratio at different wavelengths and the object temperature; the radiation emittance ratio formula of the object at a preset temperature at wavelengths 、 is determined according to the numerical relationship based on the classical Planck blackbody radiation law. , wherein is the preset temperature, respectively, the radiation emittance of the object at the wavelength , , is a second radiation constant; and , are close, , the real temperature of the object is: , wherein only is a variable; the turbine blade temperature field temperature is obtained according to the radiation emittance ratio at different wavelengths and the radiation emittance ratio formula, and the calculation formula of the turbine blade temperature field temperature is: .

[0074] Optionally, the signal acquisition and data processing system is further configured to obtain a preset numerical relationship between a charge sensitivity experimental value, a feedback capacitance experimental value of a charge quantity conversion stage, a variable voltage experimental value, and a nacelle acceleration experimental value; an expression of the preset numerical relationship is: , is the variable voltage experimental value, is the feedback capacitance experimental value of the charge quantity conversion stage, is the charge sensitivity experimental value, a is the nacelle acceleration experimental value; the nacelle acceleration is calculated based on the preset numerical relationship and the variable voltage.

[0075] The device provided by the embodiments of the present disclosure can execute the steps of the aero-engine state monitoring method provided in Embodiment 1, and thus will not be described in detail.

[0076] The aero-engine state monitoring device provided by the embodiments acquires the magnetic flux variation, the light signal, the thermal radiation signal, and the charge variation through a composite sensor; converts the magnetic flux variation into a coil output voltage, converts the light signal into a light signal voltage, converts the thermal radiation signal to obtain a thermal radiation emittance voltage signal at different wavelengths, and converts the charge variation into a variable voltage; obtains a blade tip gap measured value based on the coil output voltage and a probe-to-blade tip distance value; obtains a blade vibration displacement based on the light signal voltage according to the blade tip timing principle; performs colorimetric processing on the thermal radiation emittance voltage signal at different wavelengths to obtain a radiation emittance ratio at different wavelengths, and further obtain a turbine blade temperature field temperature; and calculates the nacelle acceleration based on the variable voltage. In this way, the composite sensor is used to simultaneously acquire the blade tip gap measured value, the blade vibration displacement, the turbine blade temperature field temperature, and the nacelle acceleration at the same measurement point, which simplifies the engineering test steps and provides a reliable monitoring method for aero-engine health management and fault prediction.

[0077] Embodiment 3 The embodiment of the present disclosure provides a computer readable storage medium, which stores a computer program, and the computer program is executed by a processor to implement the aero-engine state monitoring method in the embodiment 1.

[0078] In the embodiment, the computer readable storage medium can be a read-only memory (ROM), a random access memory (RAM), a magnetic disk or an optical disk, etc.

[0079] The computer readable storage medium provided by the embodiment can implement the aero-engine state monitoring method provided by the embodiment 1, and details are not described herein to avoid repetition.

[0080] In all the examples shown and described herein, any specific values should be interpreted as merely exemplary and not as a limitation, and thus other examples of the example embodiments can have different values.

[0081] It should be noted that similar reference numbers and letters refer to similar items throughout the drawings, and thus once an item is defined in one drawing, it need not be further defined and explained in subsequent drawings.

[0082] The above-described embodiments only express several implementation manners of the present application, and the description is relatively specific and detailed, but it should not be understood as a limitation on the scope of the present application. It should be pointed out that for those skilled in the art, several modifications and improvements can be made without departing from the concept of the present application, and these all belong to the protection scope of the present application.

Claims

1. An aeroengine condition monitoring method characterised by, The method comprises the following steps: acquiring a magnetic flux change amount, a light signal, a thermal radiation signal, and a charge change amount through a composite sensor; converting the magnetic flux change amount into a coil output voltage, converting the light signal into a light signal voltage, converting the thermal radiation signal into a thermal radiation exitance voltage signal at different wavelengths, and converting the charge change amount into a change voltage; obtaining a blade tip clearance measured value according to the coil output voltage and a probe-to-blade tip spacing value; obtaining a blade vibration displacement according to the light signal voltage based on a blade tip timing principle; performing colorimetric processing on the thermal radiation exitance voltage signal at different wavelengths to obtain a radiation exitance ratio at different wavelengths, and obtaining a turbine blade temperature field temperature according to the radiation exitance ratio at different wavelengths; calculating a casing acceleration according to the change voltage.

2. The aeroengine condition monitoring method of claim 1, wherein, The method of obtaining the blade tip clearance measured value according to the coil output voltage and the probe-to-blade tip spacing value comprises the following steps: obtaining a preset mapping relationship between a coil output voltage experimental value and a probe-to-blade tip spacing experimental value; calculating the blade tip clearance measured value according to the coil output voltage based on the preset mapping relationship.

3. The aeroengine condition monitoring method of claim 1, wherein, The method of obtaining the blade vibration displacement according to the light signal voltage based on the blade tip timing principle comprises the following steps: when a voltage amplitude of the light signal voltage reaches a preset threshold, determining a time corresponding to the current light signal voltage as a blade arrival timestamp; calculating a blade tip line speed according to a preset key phase timestamp; calculating the blade vibration displacement according to the blade tip line speed, the preset key phase timestamp, the blade arrival timestamp, and a preset arc length between a blade and a key phase.

4. The aircraft engine condition monitoring method of claim 1, wherein, The method of obtaining the turbine blade temperature field temperature according to the radiation exitance ratio at different wavelengths comprises the following steps: obtaining a numerical relationship between the radiation exitance ratio at different wavelengths and an object temperature; Based on the classical Planck blackbody radiation law, the ratio of the radiant exitance of an object at a preset temperature at wavelengths , is determined according to the numerical relationship , wherein is the preset temperature, is the radiant exitance of the object at wavelengths , , is the second radiation constant; and , when the two are close, , the real temperature of the object is: , wherein only is a variable; The turbine blade temperature field temperature is obtained according to the radiation emission degree ratio at the different wavelengths and the radiation emission degree ratio formula, and the calculation formula of the turbine blade temperature field temperature is: .

5. The aircraft engine condition monitoring method of claim 1, wherein, The method of calculating the casing acceleration according to the change voltage comprises the following steps: acquire a preset numerical relationship among an experimental value of charge sensitivity, an experimental value of feedback capacitance of a charge quantity conversion stage, an experimental value of change voltage, and an experimental value of case acceleration; an expression of the preset numerical relationship is: , is the experimental value of change voltage, is the experimental value of feedback capacitance of the charge quantity conversion stage, is the experimental value of charge sensitivity, a is the experimental value of case acceleration; calculating the casing acceleration according to the change voltage based on the preset numerical relationship.

6. An aeroengine condition monitoring apparatus characterised in that, The device for executing the aero-engine state monitoring method of claim 1 comprises a composite sensor, an eddy current preamplifier, a vibration photoelectric conversion module, a temperature photoelectric conversion module, a charge amplifier, and a signal acquisition and data processing system. The composite sensor is used to acquire a magnetic flux change amount, a light signal, a thermal radiation signal, and a charge change amount. The eddy current preamplifier is used to convert the magnetic flux change amount into a coil output voltage and output the coil output voltage to the signal acquisition and data processing system. The vibration photoelectric conversion module is used to convert the light signal into a light signal voltage and output the light signal voltage to the signal acquisition and data processing system. The temperature photoelectric conversion module is used to convert the thermal radiation signal into a thermal radiation signal at different wavelengths and convert the thermal radiation signal at different wavelengths into an exitance voltage signal output to the signal acquisition and data processing system. The charge amplifier is used to convert the charge change amount into a change voltage and output the change voltage to the signal acquisition and data processing system. The signal acquisition and data processing system is used to obtain a blade tip clearance measured value according to the coil output voltage and the probe distance from the blade tip clearance value; based on the blade tip timing principle, the blade vibration displacement is obtained according to the light signal voltage; the thermal radiation exitance voltage signals under different wavelengths are subjected to colorimetric processing to obtain the radiation exitance ratio at different wavelengths, and the turbine blade temperature field temperature is obtained according to the radiation exitance ratio at different wavelengths; and the computer case acceleration is calculated according to the change voltage.

7. The aircraft engine condition monitoring apparatus of claim 6 wherein, The composite sensor comprises a sensor probe, a signal distributor, an emitting optical fiber, an eddy current signal line, a temperature receiving optical fiber, a vibration receiving optical fiber and an acceleration signal line. The sensor probe is installed on the composite sensor case. The signal distributor is used to transmit the magnetic flux change amount to the eddy current signal line, transmit the light signal to the vibration receiving optical fiber, transmit the thermal radiation signal to the temperature receiving optical fiber, and transmit the charge change amount to the acceleration signal line.

8. The aircraft engine condition monitoring apparatus of claim 7 wherein, The sensor probe comprises a signal line transmission channel, which is connected with the emitting optical fiber, the temperature receiving optical fiber, the vibration receiving optical fiber, the eddy current signal line and the acceleration signal line.

9. The aircraft engine condition monitoring apparatus of claim 8 wherein, The sensor probe further comprises a composite sensor shell and a piezoelectric module, the piezoelectric module is a nut type piezoelectric module, and is installed on the composite sensor shell through threaded connection.

10. A computer-readable storage medium, characterized in that, The computer program is stored in the computer, and the computer program is executed by the processor to realize the aero-engine state monitoring method according to any one of claims 1 to 5.

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