A method for predicting aircraft structure vibration under specific working conditions
The method uses vibration transmissibility and cross-power spectral density matrices to accurately predict aircraft structure vibration, addressing the limitations of existing methods and improving the assessment and design of airborne equipment.
Patent Information
- Application Number
- CN202111554412.6
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- Filing Date
- 2021-12-17
- Publication Date
- 2024-12-31
- Estimated Expiration
- 2041-12-17
AI Technical Summary
Current methods for predicting aircraft structure vibration under specific working conditions, such as calculation, statistical, and empirical formula methods, lack accuracy, reliability, and are subjective, making it difficult to assess and improve airborne equipment performance.
A method involving the use of vibration transmissibility matrices and cross-power spectral density matrices, utilizing sensors to measure time domain information from easily measurable and non-measurable vibration response points, to predict aircraft structure vibration accurately.
Enables accurate prediction of aircraft structure vibration under specific conditions, enhancing the reliability of assessing and improving airborne equipment performance.
Smart Images
Figure CN114186349B_ABST
Abstract
Description
Technical Field
[0001] The present application belongs to the technical field of aircraft structure vibration prediction under specific working conditions, and specifically relates to a method for predicting aircraft structure vibration under specific working conditions. Background Art
[0002] According to statistics, about one-third of the failures of onboard equipment on an aircraft are caused by the vibration of the aircraft structure. Determining the vibration conditions of the aircraft structure under various working conditions is of great significance for the assessment, design and improvement of airborne equipment.
[0003] Under certain working conditions, the vibration of the aircraft structure can be directly obtained by measuring each vibration response point on it. However, under certain specific working conditions, the vibration of some vibration response points on the aircraft structure is difficult to measure directly. Therefore, the vibration of the aircraft structure under specific working conditions is currently predicted mainly through the following methods:
[0004] 1) Calculation method, which involves many assumptions and it is difficult to guarantee the accuracy of the results;
[0005] 2) Statistical method, which is based on the envelopment of measured data and has a simple data processing method, and there are cases where the results are quite different from the actual situation;
[0006] 3) The empirical formula method lacks rigorous derivation and is difficult to tailor or modify. It is extremely dependent on the experience of technical personnel, is greatly influenced by subjectivity, and is not easy to promote and apply.
[0007] This application is proposed in view of the above-mentioned technical defects.
[0008] It should be noted that the disclosure of the above background technology content is only used to assist in understanding the inventive concept and technical solution of the present invention, and it does not necessarily belong to the prior art of this patent application. In the absence of clear evidence that the above content has been disclosed on the filing date of this application, the above background technology should not be used to evaluate the novelty and creativity of the present application. Summary of the invention
[0009] The purpose of the present application is to provide a method for predicting aircraft structure vibration conditions under specific working conditions, so as to overcome or alleviate at least one of the known technical deficiencies.
[0010] The technical solution of this application is:
[0011] A method for predicting aircraft structure vibration under specific working conditions, comprising:
[0012] Based on the time domain information of vibration response points that are easy to directly measure under specific conditions on the aircraft structure under test conditions and the time domain information of vibration response points that are not easy to directly measure under specific conditions, the vibration transmissibility matrix under the aircraft structure test conditions is obtained:
[0013]
[0014] Based on the time domain information of vibration response points on the aircraft structure that are easy to directly measure under specific working conditions and the vibration transmissibility matrix under the aircraft structure test conditions, the time domain information of vibration response points on the aircraft structure that are not easy to directly measure under specific working conditions is calculated:
[0015]
[0016] in,
[0017] is the vibration transmissibility matrix under the aircraft structure test condition;
[0018] It is easy to directly measure the vibration response of the aircraft structure under specific conditions under test conditions.
[0019] The cross power spectrum density matrix between the vibration response points and the vibration response points that are difficult to measure directly;
[0020] It is easy to directly measure the vibration response of the aircraft structure under specific conditions under test conditions.
[0021] The cross-power spectral density matrix between the response points;
[0022] It is easy to directly measure the vibration of the aircraft structure under specific working conditions.
[0023] response points, and the cross-power spectrum density matrix between vibration response points that are not easy to measure directly;
[0024] It is easy to directly measure the vibration of the aircraft structure under specific working conditions.
[0025] Cross-power spectral density matrix between response points.
[0026] According to at least one embodiment of the present application, in the above-mentioned method for predicting aircraft structure vibration conditions under specific working conditions,
[0027]
[0028] in,
[0029] is the Kth test condition, the Kth specific condition on the aircraft structure
[0030] It is difficult to directly measure the vibration response point under the U specific working condition.
[0031] The cross power spectral density of
[0032] is the Kth test condition, the Kth specific condition on the aircraft structure
[0033] It is easy to directly measure the vibration response point under the Kth specific working condition.
[0034] The cross power spectral density.
[0035] According to at least one embodiment of the present application, in the above-mentioned method for predicting aircraft structure vibration conditions under specific working conditions,
[0036]
[0037] in,
[0038] Under specific working conditions, the aircraft structure is easy to directly
[0039] Measuring vibration response points, the mutual power between vibration response points is not easy to measure directly under the Uth specific working condition
[0040] rate spectral density;
[0041] is the Kth specific working condition on the aircraft structure that is easy to be directly
[0042] Directly measure the vibration response points, and the mutual measurement between the vibration response points under the Kth specific working condition is easy
[0043] Power spectral density.
[0044] According to at least one embodiment of the present application, in the above-mentioned method for predicting the vibration condition of the aircraft structure under specific working conditions, the time domain information of the vibration response points that are easy to directly measure under specific working conditions and the time domain information of the vibration response points that are not easy to directly measure under specific working conditions on the aircraft structure under test working conditions are measured and obtained under test working conditions by sensors arranged at the vibration response points that are easy to directly measure under specific working conditions and the vibration response points that are not easy to directly measure under specific working conditions;
[0045] The time domain information of the vibration response points on the aircraft structure under specific working conditions that are easy to directly measure under specific working conditions is measured under specific working conditions by sensors arranged at the vibration response points that are easy to directly measure under specific working conditions.
[0046] According to at least one embodiment of the present application, in the above-mentioned method for predicting aircraft structure vibration conditions under specific working conditions, the sensor is an acceleration sensor.
[0047] This application has at least the following beneficial technical effects:
[0048] A method for predicting the vibration of an aircraft structure under specific working conditions is provided. The method obtains a vibration transmissibility matrix of the aircraft structure under the test working conditions based on the time domain information of vibration response points that are easy to directly measure under the specific working conditions on the aircraft structure under the test working conditions and the time domain information of vibration response points that are not easy to directly measure under the specific working conditions. Subsequently, the time domain information of vibration response points that are not easy to directly measure under the specific working conditions on the aircraft structure is calculated based on the vibration transmissibility matrix under the test working conditions of the aircraft structure and the time domain information of vibration response points that are easy to directly measure under the specific working conditions on the aircraft structure under the specific working conditions, thereby realizing the prediction of the vibration of the aircraft structure under the specific working conditions.
[0049] The aircraft structure vibration transmissibility matrix describes the relationship between the power spectral densities of two vibration response points in a multi-degree-of-freedom system and is unique. Based on the time domain information of the vibration response points that are easy to directly measure under specific conditions on the aircraft structure under the test conditions and the time domain information of the vibration response points that are not easy to directly measure under specific conditions, the vibration transmissibility matrix under the aircraft structure test conditions is obtained. Subsequently, the vibration transmissibility matrix under the aircraft structure test conditions and the time domain information of the vibration response points that are easy to directly measure under specific conditions on the aircraft structure under specific conditions are used to calculate the time domain information of the vibration response points that are not easy to directly measure under specific conditions on the aircraft structure, so as to achieve a high reliability in predicting the vibration conditions of the aircraft structure under specific conditions. BRIEF DESCRIPTION OF THE DRAWINGS
[0050] Figure 1 is a schematic diagram of a method for predicting aircraft structure vibration conditions under specific working conditions provided in an embodiment of the present application;
[0051] Figure 2 It is a schematic diagram of the distribution of response points and excitation points on the aircraft wing structure provided in an embodiment of the present application;
[0052] Figure 3 It is a schematic diagram of the distribution of acceleration sensors on the fuselage and left and right wings of an aircraft provided in an embodiment of the present application;
[0053] Figure 4 It is a time domain acceleration diagram of the entire measurement period of the measurement point provided in the embodiment of the present application;
[0054] Figure 5 It is a schematic diagram provided in an embodiment of the present application, which establishes a vibration transmissibility matrix based on data of measuring points under two working conditions, calculates data of measuring points under a third working condition, and compares the data with the data under the third working condition.
[0055] In order to better illustrate this embodiment, some parts of the drawings may be omitted, enlarged or reduced, and do not represent the size of the actual product; in addition, the drawings are used for illustrative purposes, and their positional relationships are limited to illustrative purposes and cannot be understood as limitations on this patent. DETAILED DESCRIPTION
[0056] In order to make the technical solution and advantages of the present application clearer, the technical solution of the present application will be further described in detail in detail and in detail in conjunction with the accompanying drawings. It can be understood that the specific embodiments described here are only partial embodiments of the present application, which are only used to explain the present application, not to limit the present application. It should be noted that, for the convenience of description, only the parts related to the present application are shown in the accompanying drawings, and other related parts can refer to the general design. In the absence of conflict, the embodiments in the present application and the technical features in the embodiments can be combined with each other to obtain new embodiments.
[0057] In addition, unless otherwise defined, the technical terms or scientific terms used in the description of this application shall be the common meanings understood by the general technicians in the field to which this application belongs. The words "upper", "lower", "left", "right", "center", "vertical", "horizontal", "inside", "outside" and other words indicating orientation used in the description of this application are only used to indicate the relative direction or positional relationship, and do not imply that the device or element must have a specific orientation, be constructed and operated in a specific orientation. When the absolute position of the described object changes, its relative positional relationship may also change accordingly, so it cannot be understood as a limitation on this application. The "first", "second", "third" and similar terms used in the description of this application are only used for descriptive purposes to distinguish different components, and cannot be understood as indicating or implying relative importance. The similar words "one", "one" or "the" used in the description of this application should not be understood as an absolute limitation on quantity, but should be understood as the existence of at least one. The similar words "including" or "comprising" used in the description of this application mean that the elements or objects appearing in front of the word include the elements or objects listed after the word and their equivalents, without excluding other elements or objects.
[0058] In addition, it should be noted that, unless otherwise clearly specified and limited, the words "installed", "connected", "connected" and similar terms used in the description of this application should be understood in a broad sense. For example, the connection can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection, an indirect connection through an intermediate medium, or the internal connection of two components. Technical personnel in the field can understand their specific meanings in this application according to the specific circumstances.
[0059] The following is combined with Figures 1 to 5 , taking a typical aircraft wing structure as an example, the present application is further described in detail.
[0060] Step 1: Identify the vibration response points on the wing structure that are difficult to directly measure under specific working conditions, which may be response point 3, response point 4, and response point 5, and arrange sensors at these three points;
[0061] Step 2: Select vibration response points on the wing structure that are easy to directly measure under specific working conditions, which may be response point 1 and response point 2, and arrange sensors at these two points;
[0062] Step 3: Select the working conditions where response point 1, response point 2, response point 3, response point 4, and response point 5 are all easy to measure directly as the test working conditions, which can be working conditions 1 and working conditions 2, where working condition 1 is only excited at excitation point 1, and working condition 2 is only excited at excitation point 2. The specific working condition is working condition 3, which is excited at excitation point 1 and excitation point 3 at the same time. Response point 1, response point 2, response point 3, response point 4, response point 5, excitation point 1, excitation point 2, and excitation point 3 are distributed on the aircraft wing structure as follows: Figure 2 shown.
[0063] Step 4: Build a test platform to test and obtain the time domain information of response point 1, response point 2, response point 3, response point 4, and response point 5 in working condition 1 and working condition 2, as well as the time domain information of response point 1 and response point 2 in working condition 3;
[0064] Step 5: Use the test data to establish the vibration transmissibility matrix. The specific process is as follows:
[0065] Case 1 applies force load only at excitation point 1.
[0066] Case 2 applies force load only at excitation point 2.
[0067] Working condition 3 is to apply force loads at excitation points 1 and 2 simultaneously.
[0068]
[0069] The calculation data of response points 1, 2, 3, 4 and 5 under working conditions 1 and 2 are used to establish the vibration transmissibility matrix of the wing structure. At this time, the known conditions are:
[0070]
[0071] Under the working conditions 1 and 2, the cross power spectrum density matrix of each response point is established:
[0072]
[0073]
[0074] Assuming that working conditions 1 and 2 are represented by a and b, the vibration transmissibility matrix of the wing structure is:
[0075]
[0076] Step 6: Using the wing structure vibration transmissibility matrix established in the previous step, under working condition 3, the cross-power spectrum density matrix between response point 1 and response point 2 is used to obtain the cross-power spectrum density matrix between response point 1, response point 2, response point 3, response point 4, and response point 5, and the frequency domain information of response point 3, response point 4, and response point 5 under working condition 3 is obtained to predict the vibration of the wing structure:
[0077]
[0078] In a specific test case, the relevant loads on the outer surface of the fuselage of the test aircraft under different engine power states are measured under the following conditions:
[0079] Serial number Engine status First working condition Power 30% Second working condition Power 60% The third working condition Power 70% The fourth working condition Power 80% Fifth working condition Power 90% The sixth working condition Full power
[0080] A total of 37 measuring points are arranged in this test to test the vertical vibration of the aircraft fuselage and the left and right wings. Each measuring point is arranged with a unidirectional accelerometer. The installation position diagram is shown in the figure below. Figure 3 As shown, five acceleration sensors are evenly arranged on the upper surface of the fuselage along the center axis of the fuselage, and 16 acceleration sensors are arranged on each of the left and right wings.
[0081] The measurement frequency range is 0Hz~2000Hz, and the recording time is 30 seconds after each working condition is stable. Figure 4 It is the time domain acceleration diagram of measuring point 1 during the entire measuring period.
[0082] The acceleration test results of all measuring points are classified into working conditions, and the values after the working conditions are stable are taken. In order to facilitate the subsequent analysis, the longest time period after stabilization is taken as much as possible. Working condition 1 takes the time domain data of 0 to 23s, working condition 2 takes the time domain data of 38 to 88s, working condition 3 takes the time domain data of 95 to 145s, working condition 4 takes the time domain data of 152 to 192s, working condition 5 takes the time domain data of 203 to 243s, and working condition 6 takes the time domain data of 255 to 270s.
[0083] The vibration transmissibility matrix is established using different working condition data to verify the aircraft structure vibration prediction method under specific working conditions provided in this application:
[0084] 1) Establish a transmissibility matrix with the data of measuring points 1, 2, 3, 4 and 5 under two working conditions. Calculate the data of measuring points 3, 4 and 5 under the third working condition with the transmissibility matrix and the data of measuring points 1 and 2 under the third working condition. Compare the calculated data of measuring points 3, 4 and 5 under the third working condition with the measured data. Figure 5 As shown;
[0085] 2) The data of measuring point 1, measuring point 6, measuring point 10, measuring point 13, measuring point 15, measuring point 2, measuring point 3, measuring point 4, measuring point 5, measuring point 7, measuring point 8, measuring point 9, measuring point 11, measuring point 12, measuring point 14, and measuring point 16 under five working conditions are used to establish a transmission rate matrix. The data of measuring point 1, measuring point 6, measuring point 10, measuring point 13, and measuring point 15 under the sixth working condition and the data of measuring point 2, measuring point 3, measuring point 4, measuring point 5, measuring point 7, measuring point 8, measuring point 9, measuring point 11, measuring point 12, measuring point 14, and measuring point 16 under the sixth working condition are compared with the measured data. Figure 5 There are similar results. It can be concluded that according to the method disclosed in this application, a vibration transmissibility matrix is established to calculate the vibration response of the measuring point under specific working conditions, and the results are accurate and reliable.
[0086] The various embodiments in the specification are described in a progressive manner, and each embodiment focuses on the differences from other embodiments. The same or similar parts between the various embodiments can be referenced to each other.
[0087] So far, the technical solution of the present application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the scope of protection of the present application is obviously not limited to these specific embodiments. Without departing from the principles of the present application, those skilled in the art can make equivalent changes or substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will fall within the scope of protection of the present application.
Claims
1. A method for predicting aircraft structure vibration under specific working conditions, characterized in that: include: Based on the time domain information of vibration response points that are easy to directly measure under specific conditions on the aircraft structure under test conditions and the time domain information of vibration response points that are not easy to directly measure under specific conditions, the vibration transmissibility matrix under the aircraft structure test conditions is obtained: Based on the time domain information of vibration response points on the aircraft structure that are easy to directly measure under specific working conditions and the vibration transmissibility matrix under the aircraft structure test conditions, the time domain information of vibration response points on the aircraft structure that are not easy to directly measure under specific working conditions is calculated: in, is the vibration transmissibility matrix under the aircraft structure test condition; It is the cross-power spectrum density matrix between vibration response points that are easy to directly measure and difficult to directly measure under specific conditions on the aircraft structure under test conditions; It is a cross-power spectrum density matrix that is easy to directly measure the vibration response points under specific conditions on the aircraft structure under test conditions; It is a cross-power spectrum density matrix between vibration response points that are easy to directly measure and difficult to directly measure under specific working conditions on the aircraft structure; It is a cross-power spectral density matrix that is easy to directly measure vibration response points on the aircraft structure under specific working conditions.
2. The method for predicting aircraft structure vibration under specific working conditions according to claim 1, characterized in that: in, is the cross-power spectral density between the vibration response points on the aircraft structure that are easy to directly measure under the Kth specific working condition and difficult to directly measure under the Uth specific working condition under the Kth test working condition; Under the Kth test condition, it is easy to directly measure the vibration response point under the Kth specific condition on the aircraft structure, and it is easy to directly measure the cross-power spectral density between the vibration response points under the Kth specific condition.
3. The method for predicting aircraft structure vibration under specific working conditions according to claim 1, characterized in that: in, Under specific working conditions, the cross-power spectrum density between vibration response points on the aircraft structure that are easy to directly measure under the Kth specific working condition and difficult to directly measure under the Uth specific working condition; For specific working conditions, it is easy to directly measure the vibration response point on the aircraft structure under the Kth specific working condition, and it is easy to directly measure the cross-power spectral density between the vibration response points under the Kth specific working condition.
4. The method for predicting aircraft structure vibration under specific working conditions according to claim 1, characterized in that: The time domain information of the vibration response points that are easy to directly measure under specific working conditions and the time domain information of the vibration response points that are not easy to directly measure under specific working conditions on the aircraft structure under the test working conditions are measured under the test working conditions by sensors arranged at the vibration response points that are easy to directly measure under specific working conditions and the vibration response points that are not easy to directly measure under specific working conditions; The time domain information of the vibration response points on the aircraft structure under specific working conditions that are easy to directly measure under specific working conditions is measured under specific working conditions by sensors arranged at the vibration response points that are easy to directly measure under specific working conditions.
5. The method for predicting aircraft structure vibration under specific working conditions according to claim 4, characterized in that: The sensor is an acceleration sensor.