A method and system for analyzing compressor tip aerodynamic circumferential mode

Abstract

The present application relates to the technical field of multistage axial flow compressor design, and discloses a kind of compressor blade tip aerodynamic circumferential mode number analysis method and system, through the dynamic pressure measuring point of two groups of odd measuring points arranged circumferentially on compressor casing wall surface, the time domain signal synchronous acquisition of all dynamic pressure measuring points pulsating pressure on casing wall surface during the performance test process of compressor component under test condition is carried out, and it is guaranteed that two groups of odd measuring points have a common measuring point, to obtain the frequency domain information of each measuring point, and extract the main frequency of blade tip airflow excitation;Formulate reference measuring point, calculate the phase information between different measuring points and reference measuring point by cross spectrum analysis, finally linearly fit the linear relationship between the relative circumferential angle of measuring point angle and relative conversion phase by least square method, to identify the aerodynamic circumferential mode number of blade tip unstable flow.The present application can identify the aerodynamic circumferential mode number of rotor blade tip using fewer dynamic pressure sensors.

Description

Technical Field

[0001] This invention relates to the field of multi-stage axial compressor design technology, and discloses a method and system for analyzing the aerodynamic circumferential modal numbers of compressor blade tips. Background Technology

[0002] Aircraft place increasingly stringent demands on their power systems, and high performance has always been a relentless pursuit for aero-engines. To further reduce weight, engines are employing a large amount of lightweight new materials in their manufacturing process, and blade profiles exhibit more complex three-dimensional bending and twisting characteristics. These design trends all contribute to increased unsteady aerodynamic forces generated by unstable flow within the compressor, resulting in a wide-band multimodal characteristic of the airflow excitation frequency. This leads to higher dynamic stress levels on the rotor blades, often inducing high-cycle fatigue problems in aero-engine compressor blades.

[0003] The mechanism of flow-induced vibration of aero-engine compressor blades is complex. For synchronous blade vibration, the resonance of the rotor blade under different operating conditions can be determined by using the airflow excitation frequency obtained from dynamic pressure sensors and the resonance speed diagram. For asynchronous blade vibration, the airflow excitation frequency that induces asynchronous vibration is not synchronized with the rotational frequency. Its aerodynamic circumferential mode number is coupled with the overall bladed disk structure. Therefore, asynchronous vibration involves a multi-physics coupling problem of unsteady flow field, structural field, and acoustic field. Simply conducting dynamic tests on single, unrelated pulsating pressure, blade vibration stress, and noise cannot accurately obtain the blade failure mechanism and damage mode. In particular, the aerodynamic circumferential mode number generated by unstable flow at the blade tip is directly related to its coupling relationship with bladed disk vibration. Limited by the traditional Shannon-Nyquist sampling theorem, the measurement of higher-order circumferential mode numbers at the compressor rotor blade tip requires a large number of dynamic pressure sensors arranged circumferentially, which is difficult to achieve in actual compressor component performance testing. Summary of the Invention

[0004] The purpose of this invention is to provide a method and system for analyzing the aerodynamic circumferential modal number of compressor blade tips, which can identify the aerodynamic circumferential modal number of rotor blade tips using fewer dynamic pressure sensors.

[0005] To achieve the above-mentioned technical effects, the technical solution adopted by the present invention is as follows:

[0006] A method for analyzing the aerodynamic circumferential modal numbers at compressor blade tips, comprising: Multiple uniformly arranged first dynamic pressure measurement points are selected in the circumferential direction of the casing wall corresponding to the compressor rotor blade to be analyzed. Taking one of the first dynamic pressure measurement points as the starting point, multiple uniformly arranged second dynamic pressure measurement points are determined in the circumferential direction of the casing wall corresponding to the compressor rotor blade. The number of first dynamic pressure measurement points and the number of second dynamic pressure measurement points are unequal prime numbers. The time-domain signal of the pulsating pressure at all dynamic pressure measuring points in the circumferential direction of the casing wall during the compressor component performance test under the test conditions is collected synchronously. The frequency domain signal of the period when the blade amplitude of each dynamic pressure measuring point exceeds the preset threshold is obtained by Fourier transform method, and the main frequency under the excitation of the compressor rotor blade tip airflow in the frequency domain signal is extracted. Using the starting point position as the phase start point and reference zero point, the relative circumferential angles of all first dynamic pressure measuring points and second dynamic pressure measuring points are converted to obtain the relative circumferential angles of all dynamic pressure measuring points relative to the reference zero point. A second dynamic pressure measurement point adjacent to the reference zero point is selected in the circumferential direction of the casing as the phase information reference point. Cross-spectrum analysis is performed in the main frequency domain of airflow excitation to obtain the relative converted phase of the main frequency of all dynamic pressure measurement points relative to the main frequency of the phase information reference point. Based on the relative circumferential angle and relative phase information of all dynamic pressure measuring points, the least squares method is used to linearly fit the relative circumferential angle and relative converted phase of the first dynamic pressure measuring point, and to linearly fit the relative circumferential angle and relative converted phase of the second dynamic pressure measuring point. The slopes of the fitted lines of the first dynamic pressure measuring point and the second dynamic pressure measuring point are extracted and rounded. If the two rounded slopes are equal, the rounded slope is determined as the number of aerodynamic circumferential modes at the main frequency.

[0007] Furthermore, the product of the number of the first dynamic pressure measuring points and the number of the second dynamic pressure measuring points is greater than twice the number of rotor blades of the compressor to be analyzed, and the absolute value of the difference between the number of the first dynamic pressure measuring points and the number of the second dynamic pressure measuring points is not greater than 7.

[0008] Furthermore, the sampling frequencies of the first dynamic pressure measuring point and the second dynamic pressure measuring point are the same, and are greater than or equal to twice the passing frequency of the rotor blades.

[0009] Furthermore, the rotor blades pass through at a frequency according to... Analysis yielded, among which The passing frequency of the rotor blades. To assess the compressor speed corresponding to the operating conditions, This refers to the number of compressor rotor blades. The total inlet temperature of the compressor is measured in K under the test conditions.

[0010] To achieve the above technical effects, the present invention also provides a compressor tip aerodynamic circumferential modal number analysis system, used to implement the aforementioned compressor tip aerodynamic circumferential modal number analysis method, comprising: The measurement point determination module is used to select multiple uniformly arranged first dynamic pressure measurement points on the circumference of the casing wall corresponding to the compressor rotor blade to be analyzed, and to determine multiple uniformly arranged second dynamic pressure measurement points on the circumference of the casing wall corresponding to the compressor rotor blade, starting from one of the first dynamic pressure measurement points; the number of the first dynamic pressure measurement points and the number of the second dynamic pressure measurement points are unequal prime numbers. The data acquisition module is used to synchronously acquire the time-domain signal of the pulsating pressure of all dynamic pressure measuring points in the circumferential direction of the casing wall during the performance test of the compressor components under the test conditions. It obtains the frequency domain signal of the period when the blade amplitude of each dynamic pressure measuring point exceeds the preset threshold through the Fourier transform method, and extracts the main frequency of the compressor rotor blade tip airflow excitation in the frequency domain signal. The first analysis module is used to calculate the relative circumferential angles of all first dynamic pressure measuring points and second dynamic pressure measuring points, taking the starting position as the phase starting point and the reference zero point, respectively, to obtain the relative circumferential angles of all dynamic pressure measuring points relative to the reference zero point. The second analysis module is used to select a second dynamic pressure measurement point that is close to the position opposite to the reference zero point in the circumferential direction of the casing as the phase information reference point, and to perform cross-spectral analysis in the main frequency domain of airflow excitation to obtain the relative converted phase of the main frequency of all dynamic pressure measurement points relative to the main frequency of the phase information reference point. The data fitting module is used to perform linear fitting of the relative circumferential angle and relative converted phase of the first dynamic pressure measuring point using the least squares method, based on the relative circumferential angle and relative converted phase information of all dynamic pressure measuring points, and to perform linear fitting of the relative circumferential angle and relative converted phase of the second dynamic pressure measuring point. The modal number determination module is used to extract the slopes of the fitted lines of the first dynamic pressure measuring point and the second dynamic pressure measuring point respectively and round them. If the two slopes after rounding are equal, the rounded slope is determined as the aerodynamic circumferential modal number under the main frequency.

[0011] Furthermore, in the measurement point determination module, the product of the number of the first dynamic pressure measurement points and the number of the second dynamic pressure measurement points is greater than twice the number of compressor rotor blades to be analyzed, and the absolute value of the difference between the number of the first dynamic pressure measurement points and the number of the second dynamic pressure measurement points is not greater than 7.

[0012] Furthermore, in the data acquisition module, the sampling frequencies of the first dynamic pressure measuring point and the second dynamic pressure measuring point are the same, and are greater than or equal to twice the passing frequency of the rotor blades.

[0013] Furthermore, in the data acquisition module, the rotor blades pass through at a frequency according to... Analysis yielded, among which The passing frequency of the rotor blades. To assess the compressor speed corresponding to the operating conditions, This refers to the number of compressor rotor blades. The total inlet temperature of the compressor is measured in K under the test conditions.

[0014] Compared with the prior art, the beneficial effects of this invention are as follows: This invention uses two sets of odd-numbered dynamic pressure measuring points arranged circumferentially on the compressor casing wall to synchronously acquire the time-domain signals of the pulsating pressure of all dynamic pressure measuring points circumferentially on the casing wall during the performance test of compressor components under the test conditions, and ensures that the two sets of odd-numbered measuring points have a common measuring point to obtain the frequency domain information of each measuring point and extract the excitation frequency of the blade tip airflow; a reference measuring point is established, and the phase information between different measuring points and the reference measuring point is calculated through cross-spectral analysis. Finally, the linear relationship between the relative circumferential angle and the relative converted phase of the measuring point angle is linearly fitted using the least squares method to identify the aerodynamic circumferential mode number of unstable flow at the blade tip, laying the foundation for the induction mechanism and failure analysis of compressor blade flow-induced vibration problems. Attached Figure Description

[0015] Figure 1 This is a flowchart of the compressor blade tip aerodynamic circumferential modal number analysis method in Example 1 or 2; Figure 2 This is a block diagram of the compressor blade tip aerodynamic circumferential modal number analysis system in Example 1; Figure 3 This is a schematic diagram showing the distribution of dynamic pressure measuring points in Example 2; The module comprises: 1. Measurement point determination module; 2. Data acquisition module; 3. First analysis module; 4. Second analysis module; 5. Data fitting module; and 6. Modal number determination module. Detailed Implementation

[0016] The present invention will now be described in further detail with reference to the embodiments and accompanying drawings. However, this should not be construed as limiting the scope of the above-described subject matter of the present invention to the following embodiments; all technologies implemented based on the content of the present invention fall within the scope of the present invention.

[0017] Example 1 See Figure 1 and Figure 2 A method for analyzing the aerodynamic circumferential modal numbers of compressor blade tips, comprising: Multiple uniformly arranged first dynamic pressure measurement points are selected in the circumferential direction of the casing wall corresponding to the compressor rotor blade to be analyzed. Taking one of the first dynamic pressure measurement points as the starting point, multiple uniformly arranged second dynamic pressure measurement points are determined in the circumferential direction of the casing wall corresponding to the compressor rotor blade. The number of first dynamic pressure measurement points and the number of second dynamic pressure measurement points are unequal prime numbers. The time-domain signal of the pulsating pressure at all dynamic pressure measuring points in the circumferential direction of the casing wall during the compressor component performance test under the test conditions is collected synchronously. The frequency domain signal of the period when the blade amplitude of each dynamic pressure measuring point exceeds the preset threshold is obtained by Fourier transform method, and the main frequency under the excitation of the compressor rotor blade tip airflow in the frequency domain signal is extracted. Using the starting point position as the phase start point and reference zero point, the relative circumferential angles of all first dynamic pressure measuring points and second dynamic pressure measuring points are converted to obtain the relative circumferential angles of all dynamic pressure measuring points relative to the reference zero point. A second dynamic pressure measurement point adjacent to the reference zero point is selected in the circumferential direction of the casing as the phase information reference point. Cross-spectrum analysis is performed in the main frequency domain of airflow excitation to obtain the relative converted phase of the main frequency of all dynamic pressure measurement points relative to the main frequency of the phase information reference point. Based on the relative circumferential angle and relative phase information of all dynamic pressure measuring points, the least squares method is used to linearly fit the relative circumferential angle and relative converted phase of the first dynamic pressure measuring point, and to linearly fit the relative circumferential angle and relative converted phase of the second dynamic pressure measuring point. The slopes of the fitted lines of the first dynamic pressure measuring point and the second dynamic pressure measuring point are extracted and rounded. If the two rounded slopes are equal, the rounded slope is determined as the number of aerodynamic circumferential modes at the main frequency.

[0018] In this embodiment, two sets of odd-numbered dynamic pressure sensors are arranged circumferentially on the compressor casing wall to synchronously acquire the time-domain signals of the pulsating pressure at all dynamic pressure sensors on the casing wall during the performance test of the compressor components under evaluation conditions, ensuring that the two sets of odd-numbered sensors share a common measurement point. The time-domain information of the pulsating pressure at different circumferential measurement points during the period of maximum amplitude is selected, and the frequency-domain information of each measurement point is obtained through Fourier transform to extract the excitation frequency of the blade tip airflow. A reference measurement point is established, and the phase information between different measurement points and the reference measurement point is calculated through cross-spectral analysis. Finally, the linear relationship between the relative circumferential angle and the relative converted phase of the measurement point angle is linearly fitted using the least squares method to identify the aerodynamic circumferential mode number of the unstable flow at the blade tip. This invention can identify the aerodynamic circumferential mode number of the rotor blade tip using fewer dynamic pressure sensors, laying the foundation for the induction mechanism and failure analysis of flow-induced vibration problems in compressor blades, thereby improving the reliability and safety of aero-engine compressor blade operation.

[0019] Based on the same inventive concept, this embodiment also provides a compressor tip aerodynamic circumferential modal number analysis system for implementing the aforementioned compressor tip aerodynamic circumferential modal number analysis method, including: The measurement point determination module 1 is used to select multiple uniformly arranged first dynamic pressure measurement points on the circumference of the casing wall corresponding to the compressor rotor blade to be analyzed, and to determine multiple uniformly arranged second dynamic pressure measurement points on the circumference of the casing wall corresponding to the compressor rotor blade, starting from one of the first dynamic pressure measurement points; the number of the first dynamic pressure measurement points and the number of the second dynamic pressure measurement points are unequal prime numbers. Data acquisition module 2 is used to synchronously acquire the time domain signal of the pulsating pressure of all dynamic pressure measuring points in the circumferential direction of the casing wall during the performance test of the compressor components under the test conditions. It obtains the frequency domain signal of the period when the blade amplitude of each dynamic pressure measuring point exceeds the preset threshold through the Fourier transform method, and extracts the main frequency of the compressor rotor blade tip airflow excitation in the frequency domain signal. The first analysis module 3 is used to calculate the relative circumferential angles of all first dynamic pressure measuring points and second dynamic pressure measuring points, taking the starting position as the phase starting point and the reference zero point, respectively, to obtain the relative circumferential angles of all dynamic pressure measuring points relative to the reference zero point. The second analysis module 4 is used to select a second dynamic pressure measurement point that is close to the position opposite to the reference zero point in the circumferential direction of the casing as the phase information reference point, and to perform cross-spectral analysis in the main frequency domain of airflow excitation to obtain the relative converted phase of the main frequency of all dynamic pressure measurement points relative to the main frequency of the phase information reference point. The data fitting module 5 is used to perform linear fitting of the relative circumferential angle and relative converted phase of the first dynamic pressure measuring point angle using the least squares method based on the relative circumferential angle and relative converted phase information of all dynamic pressure measuring points, and to perform linear fitting of the relative circumferential angle and relative converted phase of the second dynamic pressure measuring point angle. The modal number determination module 6 is used to extract the slopes of the fitted lines of the first dynamic pressure measuring point and the second dynamic pressure measuring point respectively and round them. If the two slopes after rounding are equal, the rounded slope is determined as the aerodynamic circumferential mode number under the main frequency.

[0020] Example 2 See Figure 1 and Figure 3 This embodiment takes a certain type of compressor as an example, which has 21 rotor blades. The method for analyzing the aerodynamic circumferential modal number of compressor blade tips according to the present invention is described in detail, and specifically includes the following steps: Step 1: Select multiple uniformly arranged first dynamic pressure measurement points on the circumference of the casing wall corresponding to the compressor rotor blades to be analyzed. Taking one of the first dynamic pressure measurement points as the starting point, determine multiple uniformly arranged second dynamic pressure measurement points on the circumference of the casing wall corresponding to the compressor rotor blades. The number of first dynamic pressure measurement points and the number of second dynamic pressure measurement points are unequal prime numbers. To ensure accurate measurement of the compressor mode number, the product of the number of the first dynamic pressure measuring points and the number of the second dynamic pressure measuring points is greater than twice the number of rotor blades of the compressor to be analyzed, and the absolute value of the difference between the number of the first dynamic pressure measuring points and the number of the second dynamic pressure measuring points is not greater than 7.

[0021] Based on the structure of this type of compressor components, the schematic diagram of the layout of each dynamic pressure measuring point in this embodiment is as follows: Figure 3 As shown, a certain number of dynamic pressure sensors are arranged circumferentially on the casing wall at the 10% chord length position of the compressor rotor blade. The dynamic pressure sensors on the casing wall are divided into Group I (first dynamic pressure measuring point) and Group II (second dynamic pressure measuring point). Group I has 5 measuring points evenly distributed, numbered C01 to C05; Group II has 9 measuring points evenly distributed, numbered C01, C10 to C17, ensuring that Group I and Group II have a common measuring point C01, for a total of 13 measuring points circumferentially.

[0022] Step 2: Synchronously collect the time-domain signal of the pulsating pressure at all dynamic pressure measuring points in the circumferential direction of the casing wall during the compressor component performance test under the test conditions. Obtain the frequency domain signal of the period when the blade amplitude of each dynamic pressure measuring point exceeds the preset threshold by using the Fourier transform method, and extract the main frequency of the compressor rotor blade tip airflow excitation in the frequency domain signal. In this embodiment, when the dynamic pressure sensor collects data, the sampling frequencies of the first and second dynamic pressure measuring points are the same and greater than or equal to twice the rotor blade passing frequency. This ensures that the collected frequencies cover the rotor blade passing frequency; otherwise, some critical frequencies will be lost. The rotor blade passing frequency is determined according to... Analysis yielded, among which The passing frequency of the rotor blades. To assess the compressor speed corresponding to the operating conditions, This refers to the number of compressor rotor blades. The total inlet temperature of the compressor is measured in K under the test conditions.

[0023] Step 3: Using the starting point position as the phase start point and reference zero point, calculate the relative circumferential angles of all first dynamic pressure measuring points and second dynamic pressure measuring points respectively to obtain the relative circumferential angles of all dynamic pressure measuring points relative to the reference zero point.

[0024] Step 4: Select a second dynamic pressure measurement point that is close to the opposite position of the reference zero point in the circumferential direction of the casing as the phase information reference point. Perform cross-spectral analysis in the main frequency domain of airflow excitation to obtain the relative converted phase of the main frequency of all dynamic pressure measurement points relative to the main frequency of the phase information reference point.

[0025] Step 5: Based on the relative circumferential angle and relative phase information of all dynamic pressure measuring points, the least squares method is used to perform linear fitting of the relative circumferential angle and relative converted phase of the first dynamic pressure measuring point, and linear fitting of the relative circumferential angle and relative converted phase of the second dynamic pressure measuring point.

[0026] Step 6: Extract the slopes of the fitted lines of the first dynamic pressure measuring point and the second dynamic pressure measuring point respectively, and round them. If the two slopes after rounding are equal, then the rounded slope is determined as the number of aerodynamic circumferential modes under the main frequency.

[0027] In some other embodiments, if the two slopes after rounding are not equal, the time-domain signal of the pulsating pressure of all dynamic pressure measurement points in the circumferential direction of the casing wall during the compressor component performance test is divided into 3 segments. The slope after rounding is calculated for each segment, and cross-spectral analysis is performed on each segment and the corresponding straight line is fitted. The data with equal slopes after rounding are selected as the number of aerodynamic circumferential modes under the main frequency.

[0028] The above are merely preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent substitutions, and improvements made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A method for analyzing the aerodynamic circumferential modal numbers at the tip of a compressor blade, characterized in that, include: Multiple uniformly arranged first dynamic pressure measurement points are selected in the circumferential direction of the casing wall corresponding to the compressor rotor blade to be analyzed. Taking one of the first dynamic pressure measurement points as the starting point, multiple uniformly arranged second dynamic pressure measurement points are determined in the circumferential direction of the casing wall corresponding to the compressor rotor blade. The number of first dynamic pressure measurement points and the number of second dynamic pressure measurement points are unequal prime numbers. The time-domain signal of the pulsating pressure at all dynamic pressure measuring points in the circumferential direction of the casing wall during the compressor component performance test under the test conditions is collected synchronously. The frequency domain signal of the period when the blade amplitude of each dynamic pressure measuring point exceeds the preset threshold is obtained by Fourier transform method, and the main frequency under the excitation of the compressor rotor blade tip airflow in the frequency domain signal is extracted. Using the starting point position as the phase start point and reference zero point, the relative circumferential angles of all first dynamic pressure measuring points and second dynamic pressure measuring points are converted to obtain the relative circumferential angles of all dynamic pressure measuring points relative to the reference zero point. A second dynamic pressure measurement point adjacent to the reference zero point is selected in the circumferential direction of the casing as the phase information reference point. Cross-spectrum analysis is performed in the main frequency domain of airflow excitation to obtain the relative converted phase of the main frequency of all dynamic pressure measurement points relative to the main frequency of the phase information reference point. Based on the relative circumferential angle and relative phase information of all dynamic pressure measuring points, the least squares method is used to linearly fit the relative circumferential angle and relative converted phase of the first dynamic pressure measuring point, and to linearly fit the relative circumferential angle and relative converted phase of the second dynamic pressure measuring point. The slopes of the fitted lines of the first dynamic pressure measuring point and the second dynamic pressure measuring point are extracted and rounded. If the two rounded slopes are equal, the rounded slope is determined as the number of aerodynamic circumferential modes at the main frequency.

2. The method for analyzing the aerodynamic circumferential modal numbers of compressor blade tips according to claim 1, characterized in that, The product of the number of the first dynamic pressure measuring points and the number of the second dynamic pressure measuring points is greater than twice the number of rotor blades of the compressor to be analyzed, and the absolute value of the difference between the number of the first dynamic pressure measuring points and the number of the second dynamic pressure measuring points is not greater than 7.

3. The method for analyzing the aerodynamic circumferential modal numbers of compressor blade tips according to claim 1, characterized in that, The sampling frequencies of the first dynamic pressure measuring point and the second dynamic pressure measuring point are the same and are greater than or equal to twice the passing frequency of the rotor blades.

4. The method for analyzing the aerodynamic circumferential modal numbers of compressor blade tips according to claim 3, characterized in that, The frequency of rotor blade passage depends on Analysis yielded, among which The passing frequency of the rotor blades. To assess the compressor speed corresponding to the operating conditions, This refers to the number of compressor rotor blades. The total inlet temperature of the compressor is measured in K under the test conditions.

5. A compressor blade tip aerodynamic circumferential modal number analysis system, used to implement the compressor blade tip aerodynamic circumferential modal number analysis method according to claim 1, characterized in that, include: The measurement point determination module is used to select multiple uniformly arranged first dynamic pressure measurement points on the circumference of the casing wall corresponding to the compressor rotor blade to be analyzed, and to determine multiple uniformly arranged second dynamic pressure measurement points on the circumference of the casing wall corresponding to the compressor rotor blade, starting from one of the first dynamic pressure measurement points; the number of the first dynamic pressure measurement points and the number of the second dynamic pressure measurement points are unequal prime numbers. The data acquisition module is used to synchronously acquire the time-domain signal of the pulsating pressure of all dynamic pressure measuring points in the circumferential direction of the casing wall during the performance test of the compressor components under the test conditions. It obtains the frequency domain signal of the period when the blade amplitude of each dynamic pressure measuring point exceeds the preset threshold through the Fourier transform method, and extracts the main frequency of the compressor rotor blade tip airflow excitation in the frequency domain signal. The first analysis module is used to calculate the relative circumferential angles of all first dynamic pressure measuring points and second dynamic pressure measuring points, taking the starting position as the phase starting point and the reference zero point, respectively, to obtain the relative circumferential angles of all dynamic pressure measuring points relative to the reference zero point. The second analysis module is used to select a second dynamic pressure measurement point that is close to the position opposite to the reference zero point in the circumferential direction of the casing as the phase information reference point, and to perform cross-spectral analysis in the main frequency domain of airflow excitation to obtain the relative converted phase of the main frequency of all dynamic pressure measurement points relative to the main frequency of the phase information reference point. The data fitting module is used to perform linear fitting of the relative circumferential angle and relative converted phase of the first dynamic pressure measuring point using the least squares method, based on the relative circumferential angle and relative converted phase information of all dynamic pressure measuring points, and to perform linear fitting of the relative circumferential angle and relative converted phase of the second dynamic pressure measuring point. The modal number determination module is used to extract the slopes of the fitted lines of the first dynamic pressure measuring point and the second dynamic pressure measuring point respectively and round them. If the two slopes after rounding are equal, the rounded slope is determined as the aerodynamic circumferential modal number under the main frequency.

6. The compressor blade tip aerodynamic circumferential modal number analysis system according to claim 5, characterized in that, In the measurement point determination module, the product of the number of the first dynamic pressure measurement points and the number of the second dynamic pressure measurement points is greater than twice the number of compressor rotor blades to be analyzed, and the absolute value of the difference between the number of the first dynamic pressure measurement points and the number of the second dynamic pressure measurement points is not greater than 7.

7. The compressor blade tip aerodynamic circumferential modal number analysis system according to claim 5, characterized in that, In the data acquisition module, the sampling frequencies of the first dynamic pressure measuring point and the second dynamic pressure measuring point are the same and are greater than or equal to twice the passing frequency of the rotor blades.

8. The compressor blade tip aerodynamic circumferential modal number analysis system according to claim 5, characterized in that, In the data acquisition module, the rotor blades pass through at a frequency according to... Analysis yielded, among which The passing frequency of the rotor blades. To assess the compressor speed corresponding to the operating conditions, This refers to the number of compressor rotor blades. The total inlet temperature of the compressor is measured in K under the test conditions.

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