A sound fatigue load spectrum equivalent method considering temperature pre-stress

Abstract

The application provides a sound fatigue load spectrum equivalent method considering temperature prestress, comprising: determining a flight state and a time spectrum; determining a flight state and a sound load spectrum, corresponding the time proportion of each flight state with the measured sound pressure of the flight state to obtain a sound load and a time spectrum; equivalently processing the sound pressures under different flight states in the same frequency band in the sound load and the time spectrum, and superimposing the time corresponding to the equivalent sound pressure to obtain the equivalent time of the noise sound pressure action in each frequency band; constructing a finite element model of a typical position structure of an airplane, carrying out sound-heat joint response analysis of different frequency bands containing temperature prestress under the equivalent time and response analysis under pure sound pressure based on the finite element model, and obtaining dangerous point stresses under different temperatures and pure sound pressure; and correcting the dangerous point stresses under pure sound pressure based on the dangerous point stresses under different temperatures, and obtaining the equivalent sound pressure spectrum value based on the relationship between the corrected structure stress and random load.

Description

Technical Field

[0001] This invention belongs to the field of aircraft noise control technology, and specifically relates to an equivalent method for acoustic fatigue load spectrum considering temperature prestress. Background Technology

[0002] Current acoustic fatigue tests on aircraft structures often neglect the influence of static loads, considering only the effect of acoustic loads on the aircraft structure. Clearly, for aircraft structures subjected to combined static and other loads, such analysis results are unreliable and insufficient to accurately characterize the dynamic characteristics of the structure under these combined conditions. Currently, the acoustic fatigue load spectrum for aircraft structures relies primarily on experience from earlier, similar tests, lacking corresponding test design specifications. In particular, there is a lack of relevant design specifications for determining the thermoacoustic load spectrum, requiring extensive pre-test adjustments, which is not only time-consuming and labor-intensive but also inevitably causes fatigue damage to the test specimens. Summary of the Invention

[0003] The purpose of this invention is to provide an equivalent method for acoustic fatigue load spectrum considering temperature prestress, so as to solve or alleviate at least one problem in the background art.

[0004] The technical solution of this invention is: an equivalent method for acoustic fatigue load spectrum considering temperature prestress, comprising:

[0005] Step S1: Determine the typical mission profile and representative flight subjects of the aircraft, and obtain the flight status and time spectrum based on the typical mission profile and representative flight subjects.

[0006] Step S2: Determine the flight state and acoustic load spectrum, and match the time proportion of each flight state with the measured sound pressure of that flight state to obtain the acoustic load and time spectrum of the typical position structure of the aircraft.

[0007] Step S3: Equivalently convert the acoustic load of the typical position structure of the aircraft to the sound pressure of different flight states in the same frequency band in the time spectrum, and superimpose the time corresponding to the equivalent sound pressure to obtain the equivalent time of the noise sound pressure in each frequency band.

[0008] Step S4: Construct a finite element model of the typical position structure of the aircraft, and conduct acoustic-thermal joint response analysis and pure acoustic pressure analysis on the finite element model under equivalent time with temperature prestress in different frequency bands to obtain the critical point stress of the typical position structure of the aircraft at different temperatures and the critical point stress under pure acoustic pressure.

[0009] Step S5: Based on the critical point stress at different temperatures, the critical point stress under pure sound pressure is corrected to obtain the corrected structural stress. Based on the relationship between the corrected structural stress and random load, the equivalent sound pressure spectrum value is obtained.

[0010] In a preferred embodiment of the present invention, step S1, the process of obtaining the flight status and time spectrum based on typical mission profiles and representative flight subjects, is as follows:

[0011] Step S11: Count the number of times each representative flight subject appears for each aircraft within a certain time period. and the time of each flight for each representative flight subject In this context, the subscripts i and j represent the aircraft sortie number and the flight course number, respectively.

[0012] Step S12: Calculate the number of times each flight subject is represented for all the aircraft being counted. and flight time , , In the formula, L represents the total number of aircraft counted.

[0013] Step S13: Calculate the average number of times each aircraft performs each representative flight course. and the corresponding average time , , Thus, the flight status and time spectrum are obtained.

[0014] In a preferred embodiment of the present invention, step S2, which maps the time proportion of each flight state to the measured sound pressure of that flight state, is as follows: ;

[0015] in, The noise sound pressure level at the k-th typical location structure The corresponding time; The noise sound pressure level of the k-th typical location structure under the j-th representative flight subject The corresponding time; The sound pressure level of the k-th typical position structure in the representative flight subject j The percentage; N represents the number of flight subjects.

[0016] In a preferred embodiment of the present invention, step S3, which involves equivaling the acoustic load of a typical aircraft position structure to the sound pressure under different flight conditions in the same frequency band of the time spectrum, and simultaneously superimposing the corresponding times after the sound pressure equivalence to obtain the equivalent time of the noise sound pressure effect in each frequency band, includes:

[0017] Based on the double logarithmic fitting formula of the acoustic fatigue life curve of typical location structure Establish an equivalent conversion relationship between sound pressure level and lifespan. In the formula, 𝑊 is the sound pressure spectrum value, 𝑇 is the time required to reach failure, 𝑏 is the negative reciprocal of the slope of the double logarithmic acoustic fatigue line, 𝐶 is the material constant, W1 and W2 are the first and second sound pressure spectrum values ​​respectively, and T1 and T2 are the time required for the structure to reach failure under the first sound pressure spectrum value and the second sound pressure spectrum value respectively.

[0018] Based on the equivalent conversion relationship between sound pressure and lifetime, the equivalent time is obtained by summarizing the duration of sound pressure effects for various noise levels within the same frequency band. In the formula, W ki W kj For different noise sound pressure levels i and j experienced by the typical structural position k of the aircraft;

[0019] By superimposing the time corresponding to the sound pressure effect, the equivalent time of the noise sound pressure effect in each frequency band can be obtained: In the formula, m is the total number of typical noise states of the aircraft.

[0020] In a preferred embodiment of the present invention, step S4, which involves conducting acoustic-thermal joint response analysis on the finite element model at different frequency bands with temperature prestress under equivalent time, to obtain the critical point stress of the typical aircraft structure at different temperatures, is as follows:

[0021] a) Establish a finite element model of the typical positional structure of the aircraft;

[0022] b) Based on the material properties of typical aircraft structures at a finite number of different temperatures, temperature interpolation is performed to obtain the material properties at any temperature;

[0023] c) Perform thermal modal analysis of typical aircraft structures based on material properties at arbitrary temperatures to obtain thermal stress in typical aircraft structures.

[0024] d) Obtain the frequency response function of a typical aircraft positional structure;

[0025] e) Applying a load spectrum to the typical aircraft structure with acoustic and vibration power spectral density considering static load;

[0026] f) Perform dynamic response analysis of typical aircraft structures under thermoacoustic vibration conditions taking into account static load, thereby obtaining the stress at critical points of typical aircraft structures.

[0027] In a preferred embodiment of the present invention, step S5, which involves correcting the critical point stress under pure sound pressure based on the critical point stress at different temperatures to obtain the corrected structural stress, and then obtaining the equivalent sound pressure spectrum value based on the relationship between the corrected structural stress and the random load, is as follows:

[0028] The vibration stress amplitude under static load is equivalent to a state where the average stress is 0, according to the stress averaging correction formula. The stress averaging formula is corrected using the Goodman formula. ;

[0029] In the formula, S u S represents the material's ultimate strength. m The reference stress level is the critical point stress at different temperatures; S ai This refers to the actual stress, i.e., the stress at the critical point under pure sound pressure; S aoi This represents the corrected structural stress.

[0030] According to the theory of linear random vibration, the root mean square (RMS) value of the sound pressure random load excitation is proportional to the RMS value of the structural stress response:

[0031] In the formula, S a1 S a2 The stress amplitude is the average stress of 0; P1 and P2 are the sound pressure levels.

[0032] The response results for each state are compared with the results for the reference state to obtain the equivalent sound pressure spectrum value.

[0033] The present invention provides an equivalent method for acoustic fatigue load spectrum considering temperature prestress. By combining numerical and theoretical approaches, the method uses numerical methods to calculate the random acoustic response under prestress when considering temperature prestress, and combines the dynamic fatigue life curve of the material to perform load spectrum equivalence. This method can effectively simplify the traditional test process and save project costs. Attached Figure Description

[0034] To more clearly illustrate the technical solution provided by the present invention, a brief description of the accompanying drawings will be provided below. Obviously, the drawings described below are merely some embodiments of the present invention.

[0035] Figure 1 This is a schematic diagram of the equivalent method of acoustic fatigue load spectrum considering temperature prestress in this invention.

[0036] Figure 2 This is a schematic diagram illustrating the effect of typical thermal prestress on the acoustic fatigue life of a structure according to an embodiment of the present invention.

[0037] Figure 3 This is a schematic diagram illustrating the equivalent summarization of the noise spectrum in the same frequency band for a typical structure according to an embodiment of the present invention. Detailed Implementation

[0038] To make the objectives, technical solutions, and advantages of this invention clearer, the technical solutions in the embodiments of this invention will be described in more detail below with reference to the accompanying drawings.

[0039] To address the problem of aircraft structures (such as aircraft panels) simultaneously facing thermal and acoustic loads, with loads varying over time and lacking corresponding test specifications or load spectra, this invention proposes an equivalent method for acoustic fatigue load spectra considering temperature prestress. This method can be used in tests that comprehensively verify the thermoacoustic fatigue characteristics of aircraft structures, simplifying tests that require simultaneous temperature and acoustic load assessments to acoustic load tests alone. This invention combines numerical and theoretical approaches. When considering temperature prestress, it uses numerical methods to calculate the random acoustic response under temperature prestress and combines this with the dynamic fatigue life curves of the aircraft structural materials to perform equivalent acoustic load spectra. This effectively simplifies traditional testing procedures and saves project costs.

[0040] like Figure 1 As shown, the present invention provides an equivalent method for acoustic fatigue load spectrum considering temperature prestress, comprising the following steps:

[0041] Step S1: Determine the typical mission profile and representative flight subjects of the aircraft, and obtain the flight status and time spectrum based on the typical mission profile and representative flight subjects.

[0042] A typical mission profile of an aircraft is a trajectory that describes the changes in its flight status (such as altitude, speed, weight, configuration, etc.) over time or distance during the execution of a specific mission. It can be a graph (such as an altitude-time curve) or a detailed table of state parameters.

[0043] The flight status and time spectrum of the aircraft fleet are summarized and statistically analyzed. Each typical mission profile includes at least one representative flight course. Noise load data of the aircraft structure are obtained by conducting actual flights of these representative flight courses. To ensure the representativeness of the selected typical mission profiles and to guarantee that the merged typical mission profiles of the aircraft reflect the average usage of the aircraft during its service life, the specific principles for merging flight courses are as follows:

[0044] 1) The nature and content of the missions performed by each flight subject in the same mission profile are basically the same;

[0045] 2) The flight altitude, speed, G-force variation range and sequence of each flight subject in the same mission profile are basically similar;

[0046] 3) The flight times for each flight course within the same mission profile are roughly the same;

[0047] 4) Within the same mission profile, select flight subjects with specified parameters that are slightly above average as representative flight subjects of that mission profile. Specified parameters may include, for example, throttle lever angle PLA, rotor speed, exhaust temperature, etc. Slightly above average usually means that the parameters are in the 70%~90% working state. For example, rotor speed N1≈70% is a slightly above average state.

[0048] Based on the above process, typical mission profiles and their representative flight courses can be determined. By statistically analyzing the representative flight courses within the typical mission profiles, flight status and time spectrum can be obtained, as follows:

[0049] Step S11: Count the number of times each representative flight subject appears for each aircraft within a certain time period. and the time of each flight for each representative flight subject In this context, the subscripts i and j represent the aircraft sortie number and the flight course number, respectively.

[0050] Step S12: Calculate the number of times each flight subject is represented for all the aircraft being counted. and flight time , , In the formula, L represents the total number of aircraft counted.

[0051] Step S13: Calculate the average number of times each aircraft performs each representative flight course. and the corresponding average time , , .

[0052] Step S2: Determine the flight state and acoustic load spectrum, and match the time proportion of each flight state with the measured sound pressure of that flight state to obtain the acoustic load and time spectrum of the typical position structure of the aircraft.

[0053] For each representative flight maneuver, by combining computational fluid dynamics simulation data and measured noise load data, the noise load sound pressure level of typical aircraft structural positions can be determined. The corresponding time (Subscript k indicates the serial number of the typical position structure), thus obtaining the flight state and acoustic load spectrum.

[0054] By combining flight state with time spectrum and flight state with acoustic load spectrum, that is, by mapping the time proportion of each flight state to the measured sound pressure level of that flight state, the sound pressure level and time spectrum of the typical position structure of the aircraft can be obtained, i.e., the sound pressure level SPL. k The corresponding time ;

[0055] In the formula, It represents the sound pressure level in flight subject j. The percentage is N, which represents the total number of flight subjects.

[0056] This invention uses three typical operating conditions (representing flight maneuvers) as examples, namely, operating condition 1, operating condition 2, and operating condition 3. The temperature loads for the three typical operating conditions are set to -60℃, 20℃, and 100℃, respectively, with 20℃ as the reference temperature, i.e., operating condition 2 is the reference operating condition. The sound pressure level (in dB) for these three typical operating conditions is shown in Table 1.

[0057] Table 1 Sound pressure levels under three typical operating conditions

[0058] Center frequency (Hz) 16 31.5 63 125 250 500 1000 2000 Operating condition 1 (dB) 144.68 144.22 143.07 142.48 144.88 143.38 146.42 144.77 Operating condition 2 (dB) 145.55 142.41 148.45 145.56 146.83 145.43 147.87 141.58 Operating condition 3 (dB) 143.62 145.50 143.66 143.71 145.01 148.42 145.54 144.72

[0059] The above method can be used to obtain the noise sound pressure level of the typical aircraft position structure k under comprehensive statistical conditions. Corresponding duration For example, in this embodiment of the present invention, based on the above calculation and analysis method, the action time of a typical position structure of an aircraft under three typical operating conditions can be obtained as follows: 10s 20s 30s.

[0060] Step S3: Equivalently compare the acoustic load of the typical position structure of the aircraft with the sound pressure of the same frequency band under different flight conditions in the time spectrum, and then superimpose the time corresponding to the equivalent sound pressure to obtain the equivalent time of the noise sound pressure in each frequency band.

[0061] The acoustic fatigue life curves of typical aircraft structural positions are similar to those of conventional fatigue life curves, and can be fitted using a least-squares method with a straight line from a log-log plot. The formula for the log-log fitting of the acoustic fatigue life curves of typical aircraft structural positions is as follows: , where 𝑊 is the sound pressure spectrum value, 𝑇 is the time required to reach failure, 𝑏 is the negative reciprocal of the slope of the double logarithmic acoustic fatigue line, and 𝐶 is the material constant.

[0062] Based on the above fitting formula, an equivalent conversion relationship between sound pressure level and lifetime can be established using different sound pressure level spectra and different failure times: In the formula, W1 and W2 are the first and second acoustic pressure spectrum values, respectively, and T1 and T2 are the time required for the structure to reach destruction under the first and second acoustic pressure spectrum values, respectively.

[0063] Based on the above, the duration of sound pressure levels for various noises within the same frequency band is summarized to obtain the equivalent time: In the formula, W ki W kj For different noise sound pressure levels i and j experienced by the typical structural position k of the aircraft;

[0064] By superimposing the time corresponding to the sound pressure effect, the equivalent time of the noise sound pressure effect in each frequency band can be obtained: In the formula, m represents the total number of typical noise states of the aircraft. In this embodiment of the invention, m is taken as 3.

[0065] Step S4: Construct a finite element model of the typical position structure of the aircraft, and conduct acoustic-thermal joint response analysis and response analysis under pure sound pressure at different frequency bands with temperature prestress under equivalent time on the finite element model to obtain the critical point stress of the typical position structure of the aircraft at different temperatures and the critical point stress under pure sound pressure.

[0066] In this invention, the finite element method is used to perform acoustic and thermal response analysis on the typical positional structure of an aircraft under temperature prestress conditions in different frequency bands, and at the same time, response analysis is performed under pure sound pressure.

[0067] When performing acoustic-thermal co-simulation, the nonlinear module of a commercial finite element method can be used for calculation. First, thermal modal analysis is performed on the typical structural position of the aircraft. The calculation steps are as follows:

[0068] a) Establish a finite element model of the typical positional structure of the aircraft;

[0069] b) Based on the material properties of typical aircraft structures at a finite number of different temperatures, temperature interpolation is performed to obtain the material properties at any temperature;

[0070] c) Perform thermal modal analysis of typical aircraft structures at arbitrary temperatures based on material properties to obtain thermal stress in typical aircraft structures.

[0071] The effect of temperature on the mechanical properties of materials is mainly manifested in the following ways: as the temperature increases, the elastic modulus of the material decreases, leading to a decrease in material stiffness. Within a certain temperature range... If the elastic modulus of a material is known at certain temperatures, then the elastic modulus at any temperature t can be obtained through linear interpolation. By performing linear interpolation over each temperature interval, the elastic modulus at temperature t can be obtained. for: In the formula, This refers to the elastic modulus of the material at room temperature. This represents the change in elastic modulus due to temperature variations. For linear elastic materials, the elemental elastic matrix consists of the elastic modulus E and Poisson's ratio. Therefore, the element elastic matrix at temperature t can be expressed as the sum of the elastic matrix at the initial temperature and the additional matrix caused by the temperature difference. From this, the element stiffness matrix K at temperature t can be obtained. e (t):

[0072] ;

[0073] In the formula, Be D is a geometric matrix; e (t0) is the material constitutive matrix at the initial reference temperature t0; ΔD e O represents the change in the material constitutive matrix caused by the temperature change from t0 to t; e This is the integration domain.

[0074] Based on this, and using structural finite element theory, the dynamic equations for considering both effects can be obtained. Therefore, the structural finite element dynamic equations considering thermal effects are:

[0075] ;

[0076] In the formula, , and These represent the stiffness, damping, and mass matrix of the structure at room temperature. The displacement of the i-th node of the structure; This is the stiffness variation matrix caused by temperature. This is the stiffness variation matrix caused by stress. This is the external load excitation vector; ω is the angular frequency; j is the imaginary unit.

[0077] Subsequently, based on the structural thermal modal analysis, the thermal stress of the typical aircraft structure was added as prestress to the structural finite element model for further dynamic response analysis. The steps included:

[0078] d) Obtain the frequency response function of a typical aircraft positional structure;

[0079] e) Applying a load spectrum to the typical aircraft structure with acoustic and vibration power spectral density considering static load;

[0080] f) Perform dynamic response analysis of typical aircraft structures under thermoacoustic vibration conditions taking into account static load, thereby obtaining the stress at critical points of typical aircraft structures.

[0081] For acoustic-vibration coupling, dynamic response calculations are performed by dividing the acoustic-structure interaction mesh using finite element software. Based on the boundary conditions, the equivalent integral weak form of the sound field can be obtained as follows:

[0082] ;

[0083] In the formula, n represents the direction of the outward normal to the wall. The acoustic impedance of the sound-absorbing wall surface. Let V be the component of the vibrational acceleration of a particle on the wall along the normal direction of the wall, and V be the volume region. For a rigid wall region, For the sound-absorbing wall surface area, For an elastic wall region, This is the variation of sound pressure.

[0084] In this embodiment of the invention, a typical wall panel structure is used as an example to calculate the structural acoustic response with temperature prestress. The size of the finite element model of the wall panel is 655mm×380mm. In order to improve the critical buckling temperature of the wall panel structure, the wall panel thickness is set to 15mm and the material is 7N01-T4 aviation aluminum alloy. Its elastic modulus changes with temperature and its basic performance parameters are shown in Tables 2 and 3 below.

[0085] Table 2 Elastic Modulus of 7N01-T4 Aviation Aluminum Alloy

[0086] Temperature (°C) -60 20 100 200 Elastic modulus (GPa) 85.6 69 64.9 60.87

[0087] Table 3 Performance parameters of 7N01-T4 aerospace aluminum alloy

[0088] Poisson's ratio Density (kg / m³) coefficient of expansion Tensile strength (MPa) 0.33 2700 <![CDATA[2.4×10⁻ 5 ]]> 426

[0089] The wall panel structure is fixed on all four sides. The first critical buckling temperature of the wall panel structure is calculated to be 158.13℃. The natural frequencies of the structure under three working conditions are shown in Table 4.

[0090] Table 4. Natural frequencies of the structure under three operating conditions

[0091] First order Second order Third order Fourth level Fifth Order Sixth Order Seventh Rank Operating Condition 1 868.61 1189.2 1708.1 1976.7 Operating Condition 2 645.51 906.11 1356.2 1613.8 1853.7 1980.8 Operating Condition 3 450.05 677.08 1106.9 1379.6 1605.5 1714.2 1998.3

[0092] In acoustic response analysis, sound pressure is the pressure perpendicular to the plate. Static pressure in the same direction as the sound pressure is defined as positive pressure, and static pressure in the opposite direction to the sound pressure is defined as negative pressure.

[0093] The location of the danger point under three working conditions was determined by the root mean square value of the stress in the wall panel structure. The dynamic stress response and acoustic response data of the danger point were extracted and are shown in Table 5.

[0094] Table 5 Stress Results at Critical Points

[0095] Static response (MPa) Dynamic response (MPa) Operating Condition 1 245.301 0.997755 Operating Condition 2 0 1.31112 Operating Condition 3 186.068 2.13622

[0096] Compiling the equivalent load spectrum for a single working condition requires dynamic response analysis under reference conditions. The temperatures of both working condition 1 and working condition 3 are set as reference temperatures, and the acoustic pressure spectrum data remain unchanged as reference loads for working condition 1 and working condition 3. The response results at the critical point are shown in Table 6.

[0097] Table 6 Stress results at critical points under reference conditions

[0098] Static response (MPa) Dynamic response (MPa) Operating Condition 1 (Reference) 0 0.999956 Operating Condition 3 (Reference) 0 1.39186

[0099] Step S5: Based on the critical point stress at different temperatures, the critical point stress under pure sound pressure is corrected to obtain the corrected structural stress. Based on the relationship between the corrected structural stress and random load, the equivalent sound pressure spectrum value is obtained.

[0100] Assume the vibration stress under pure acoustic load is Sa, the duration is T, the frequency is f, and the number of cycles is N. Using the stress averaging correction formula, the vibration stress amplitude under static load is equivalent to a state where the average stress is 0. Typically, the Goodman formula can be used for correction.

[0101]

[0102] In the formula, S u S represents the material's ultimate strength. m The reference stress level is the critical point stress at different temperatures; S ai This refers to the actual stress, i.e., the stress at the critical point under pure sound pressure; S aoi This represents the corrected structural stress.

[0103] According to the theory of linear random vibration, the root mean square (RMS) value of random load excitation (such as sound pressure) is proportional to the RMS value of the structural stress response. Therefore, the relationship between the RMS value of the structural stress response and the RMS value of the sound pressure input can be expressed as:

[0104]

[0105] in , The stress amplitude when the average stress is 0; , The sound pressure level is denoted as . By comparing the response results for each state with those for the reference state, the equivalent sound pressure spectrum can be obtained.

[0106] In this embodiment of the invention, to verify the accuracy of the inductive analysis method in subsequent steps, stress equivalence is performed based on the life calculation results for each working condition. The calculation formula is as follows:

[0107]

[0108] in, For reference working conditions, dynamic stress and The fatigue life calculation results for the structure are shown in Table 7 below, representing the results under reference and normal operating conditions. The calculations for each of the three operating conditions were performed separately.

[0109] Table 7. Calculation of fatigue life and equivalent stress under three working conditions

[0110] Static stress (MPa) Dynamic stress (MPa) Fatigue life (s) Equivalent stress (MPa) Operating Condition 1 245.301 0.997755 1.2882E+19 2.00814 Operating Condition 2 0 1.31112 1.1809E+21 1.31112 Operating Condition 3 186.068 2.13622 2.7624E+15 4.47808

[0111] The fatigue characteristic curves of 7N01-T4 aerospace aluminum alloy material under three temperature conditions are as follows:

[0112] 20℃: ;

[0113] 100℃: ;

[0114] -60℃: .

[0115] Its typical impacts include Figure 2 As shown, the fatigue life of three working conditions is calculated using the Dirlik method, and then the equivalent stress is calculated using the above formula. Based on the spectral shape of the sound pressure load spectrum, the equivalent pure noise load spectrum for a single working condition can be compiled and converted into PSD (power spectral density) form for calculation and verification, as shown in Table 8.

[0116] Table 8. PSD data of equivalent noise load spectrum

[0117]

[0118] The response results and structural fatigue life calculation results are shown in Table 9.

[0119] Table 9 Equivalent Load Spectrum Response and Lifetime Calculation

[0120]

[0121] As can be seen from the table above, the error is minimal when using the structural fatigue life results of a certain working condition and its reference state for equivalence.

[0122] After obtaining the equivalent noise load spectra for the three operating conditions, the load spectrum can be summarized. First, the sound pressure spectrum values ​​within the same frequency band are summarized. According to the summarization method, a certain sound pressure level within the same frequency band is selected as the benchmark, and other sound pressure values ​​are equivalent to it to obtain the equivalent action time. Depending on the sound pressure selection criteria, it can be divided into maximum, minimum, mean, and median sound pressure levels. The results after equivalent calculations are as follows: Figure 3 As shown.

[0123] This invention addresses the urgent need for a load spectrum that can comprehensively verify the thermoacoustic fatigue characteristics of aircraft structures, given that some structures in aircraft simultaneously face thermal and acoustic loads that vary over time. This invention provides an equivalent method for acoustic fatigue load spectra considering temperature prestress. By combining numerical and theoretical methods, numerical methods are used to calculate the random acoustic response under prestress when considering temperature prestress. The load spectrum is then equivalently calculated by combining the dynamic fatigue life curve of the material. This method effectively simplifies traditional testing procedures and saves project costs.

[0124] The above description is merely a specific embodiment of the present invention, but the scope of protection of the present invention is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in the present invention should be included within the scope of protection of the present invention. Therefore, the scope of protection of the present invention should be determined by the scope of the claims.

Claims

1. An equivalent method for acoustic fatigue load spectrum considering temperature prestress, characterized in that, include: Step S1: Determine the typical mission profile and representative flight subjects of the aircraft, and obtain the flight status and time spectrum based on the typical mission profile and representative flight subjects. Step S2: Determine the flight state and acoustic load spectrum, and match the time proportion of each flight state with the measured sound pressure of that flight state to obtain the acoustic load and time spectrum of the typical position structure of the aircraft. Step S3: Equivalently convert the acoustic load of the typical position structure of the aircraft to the sound pressure of different flight states in the same frequency band in the time spectrum, and superimpose the time corresponding to the equivalent sound pressure to obtain the equivalent time of the noise sound pressure in each frequency band. Step S4: Construct a finite element model of the typical position structure of the aircraft, and conduct acoustic-thermal joint response analysis and pure acoustic pressure analysis on the finite element model under equivalent time with temperature prestress in different frequency bands to obtain the critical point stress of the typical position structure of the aircraft at different temperatures and the critical point stress under pure acoustic pressure. Step S5: Based on the critical point stress at different temperatures, the critical point stress under pure sound pressure is corrected to obtain the corrected structural stress. Based on the relationship between the corrected structural stress and random load, the equivalent sound pressure spectrum value is obtained.

2. The equivalent method for acoustic fatigue load spectrum considering temperature prestress as described in claim 1, characterized in that, In step S1, the process of obtaining the flight status and time spectrum based on typical mission profiles and representative flight subjects is as follows: Step S11: Count the number of times each representative flight subject appears for each aircraft within a certain time period. and the time of each flight for each representative flight subject In this context, the subscripts i and j represent the aircraft sortie number and the flight course number, respectively. Step S12: Calculate the number of times each flight subject is represented for all the aircraft being counted. and flight time , , In the formula, L represents the total number of aircraft counted. Step S13: Calculate the average number of times each aircraft performs each representative flight course. and the corresponding average time , , Thus, the flight status and time spectrum are obtained.

3. The equivalent method for acoustic fatigue load spectrum considering temperature prestress as described in claim 2, characterized in that, Step S2, the method for correlating the time proportion of each flight state with the measured sound pressure of that flight state is as follows: ; in, The noise sound pressure level at the k-th typical location structure The corresponding time; The noise sound pressure level of the k-th typical location structure under the j-th representative flight subject The corresponding time; The sound pressure level of the k-th typical position structure in the representative flight subject j The percentage; N represents the number of flight subjects.

4. The equivalent method for acoustic fatigue load spectrum considering temperature prestress as described in claim 3, characterized in that, Step S3, which involves equivaling the acoustic load of a typical aircraft structure with the sound pressure of different flight states in the same frequency band of the time spectrum, and then superimposing the corresponding times after the sound pressure equivalence to obtain the equivalent time of the noise sound pressure effect in each frequency band, includes: Based on the double logarithmic fitting formula of the acoustic fatigue life curve of typical location structure Establish an equivalent conversion relationship between sound pressure level and lifespan. In the formula, 𝑊 is the sound pressure spectrum value, 𝑇 is the time required to reach failure, 𝑏 is the negative reciprocal of the slope of the double logarithmic acoustic fatigue line, 𝐶 is the material constant, W1 and W2 are the first and second sound pressure spectrum values ​​respectively, and T1 and T2 are the time required for the structure to reach failure under the first sound pressure spectrum value and the second sound pressure spectrum value respectively. Based on the equivalent conversion relationship between sound pressure and lifetime, the equivalent time is obtained by summarizing the duration of sound pressure effects for various noise levels within the same frequency band. In the formula, Wki and Wkj are the different noise sound pressure levels i and j that the aircraft's typical position structure k is subjected to; By superimposing the time corresponding to the sound pressure effect, the equivalent time of the noise sound pressure effect in each frequency band can be obtained: In the formula, m is the total number of typical noise states of the aircraft.

5. The equivalent method for acoustic fatigue load spectrum considering temperature prestress as described in claim 4, characterized in that, In step S4, the process of conducting acoustic-thermal joint response analysis on the finite element model at different frequency bands with temperature prestress under equivalent time, and obtaining the critical point stress of the typical aircraft structure at different temperatures, is as follows: a) Establish a finite element model of the typical positional structure of the aircraft; b) Based on the material properties of typical aircraft structures at a finite number of different temperatures, temperature interpolation is performed to obtain the material properties at any temperature; c) Perform thermal modal analysis of typical aircraft structures based on material properties at arbitrary temperatures to obtain thermal stress in typical aircraft structures. d) Obtain the frequency response function of a typical aircraft positional structure; e) Applying a load spectrum to the typical aircraft structure with acoustic and vibration power spectral density considering static load; f) Perform dynamic response analysis of typical aircraft structures under thermoacoustic vibration conditions taking into account static load, thereby obtaining the stress at critical points of typical aircraft structures.

6. The equivalent method for acoustic fatigue load spectrum considering temperature prestress as described in claim 5, characterized in that, Step S5, the process of correcting the critical point stress under pure sound pressure based on the critical point stress at different temperatures to obtain the corrected structural stress, and obtaining the equivalent sound pressure spectrum value based on the relationship between the corrected structural stress and random load is as follows: The vibration stress amplitude under static load is equivalent to a state where the average stress is 0, according to the stress averaging correction formula. The stress averaging formula is corrected using the Goodman formula. ; In the formula, S u S represents the material's ultimate strength. m The reference stress level is the critical point stress at different temperatures; S ai This refers to the actual stress, i.e., the stress at the critical point under pure sound pressure; S aoi For the corrected structural stress, According to the theory of linear random vibration, the root mean square (RMS) value of the sound pressure random load excitation is proportional to the RMS value of the structural stress response: , In the formula, S a1 S a2 The stress amplitude is the average stress of 0; P1 and P2 are the sound pressure levels. The response results for each state are compared with the results for the reference state to obtain the equivalent sound pressure spectrum value.

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