Aircraft landing gear structure design method, system and application based on fatigue damage

Abstract

The invention relates to the technical field of aircraft engine landing gear design, and discloses a method, system and application of aircraft landing gear structure design based on fatigue damage. The invention performs fatigue life analysis on a detachment part to obtain a first limit load analysis value that takes into account fatigue damage of the detachment part under alternating stress cycles; obtains a second limit load analysis value of the detachment part based on historical data analysis of flights of aircraft of the same type, and the second limit load analysis value takes into account the combination conditions of the limit horizontal component velocity and the limit sinking component velocity under about 99% or more of normal landing conditions; selects a design shear load upper limit value of the detachment part from the first limit load analysis value and the second load analysis value, and while ensuring that the fatigue life of the detachment part meets the design requirements, also avoids the problem of a low shear load upper limit value of the detachment part, and further ensures that the detachment part of the aircraft to be designed is not prone to non-overload detachment during normal landing.

Description

Technical Field

[0001] The invention relates to the technical field of aircraft engine landing gear design, and discloses a fatigue damage-based aircraft landing gear structure design method, system and application, so as to simultaneously improve the blade strengthening effect and reduce the blade deformation. Background Art

[0002] In the design of civil aircraft, in order to minimize the damage to the surrounding wing and fuselage structure, damage to the wing fuel tank, or fire and explosion of the wing fuel tank caused by failure of the landing gear due to impact or collision, mature aircraft models often use an emergency separation structure design at the connection between the main landing gear structure and the wing. The design of the emergency separation structure can achieve safe separation of the landing gear and the wing structure when the landing gear structure is subjected to a large overload.

[0003] Traditionally, when designing emergency breakaway structures, the aircraft limited sinking speed (3.05m / s) specified in the CCAR25 airworthiness standard for transport aircraft is used as the boundary condition to analyze the load conditions of the landing gear breakaway parts when the aircraft lands under the limited sinking speed conditions, so as to design the ultimate load of the emergency breakaway structure. However, this traditional design method neither considers the horizontal component of the aircraft's landing speed nor the fatigue damage of the landing gear emergency breakaway structure during service, which leads to the problem that the breakaway parts of the aircraft are easily broken away due to fatigue damage or weak bearing capacity of the breakaway parts during normal landing. Summary of the invention

[0004] The purpose of the present invention is to provide a method, system and application for designing an aircraft landing gear structure based on fatigue damage, which not only ensures that the fatigue life of the separation part meets the design requirements, but also avoids the problem of a low upper limit of the shear load at the separation part, and further ensures that the separation part is not prone to non-overload separation during normal landing of the aircraft to be designed.

[0005] In order to achieve the above technical effects, the technical solution adopted by the present invention is:

[0006] The structural design method of aircraft landing gear based on fatigue damage includes:

[0007] Acquire a historical flight data set of the same type of aircraft, wherein the data set includes horizontal component velocity data and sinking component velocity data of the aircraft during a complete landing;

[0008] According to the size structure of the model aircraft, a three-dimensional finite element analysis model of the model aircraft is constructed, and different horizontal component velocity data and different sinking component velocity data are used as combined boundary conditions, and the three-dimensional finite element analysis model is used to analyze and obtain the stress spectrum of the landing gear separation part of the model aircraft during the process from taxiing takeoff to landing under each combined boundary condition; the separation part includes the front trunnion pin, the rear trunnion pin and the side support rod axle pin of the landing gear;

[0009] According to the maximum stress in each stress spectrum, the stress cycle combination of maximum stress-0-maximum stress is statistically extracted by using the rain flow counting method; fatigue life analysis is carried out on the three-dimensional finite element analysis model using each stress cycle combination to obtain the maximum shear strain and cycle number of the corresponding severance position under each stress cycle combination, and a strain fatigue analysis model related to the maximum shear strain and cycle number is constructed based on the Manson-Coffin formula;

[0010] According to the strain fatigue analysis model, the shear strain versus shear stress curve of the standard rod of the material at the separation site, and the number of design cycles in each separation site of the aircraft to be designed, the first limit load analysis value of each separation site of the aircraft to be designed is obtained by analysis;

[0011] A statistical method is used to obtain the maximum value of the horizontal component velocity within the range of ±3 standard deviations in the horizontal component velocity data distribution, and the maximum value of the sinking component velocity within the range of ±3 standard deviations in the sinking component velocity data distribution; and a second limit load analysis value of each separation position of the landing gear with the maximum value of the horizontal component velocity and the maximum value of the sinking component velocity as combined boundary conditions is extracted from the stress spectrum;

[0012] According to the required separation order of each separation part, the design shear load upper limit value of the corresponding separation part is selected within the interval formed by the second limit load analysis value and the first limit load analysis value of the corresponding separation part, the separation part structure design is carried out, and the aircraft landing gear structure design is completed.

[0013] Furthermore, the strain fatigue analysis model related to the maximum shear strain and the number of cycles constructed based on the Manson-Coffin formula is: ,in The maximum stress of the standard rod of the material at the separation site The shear strain amplitude under is the fatigue strength coefficient of the material at the detached part, is the shear modulus of the material at the detachment site, is the corresponding number of cycles, The fatigue ductility coefficient of the material corresponding to the separation part, , is the exponent obtained by fitting the data.

[0014] Furthermore, the method for obtaining the first limit load analysis value of each breakaway portion of the aircraft to be designed includes:

[0015] The design cycle number in the corresponding severance part of the aircraft to be designed is substituted into the strain fatigue analysis model, and the first limit load analysis value of the corresponding severance part is obtained by jointly solving the shear strain versus shear stress curve of the standard bar material of the severance part.

[0016] Furthermore, the required disconnection order of each disconnection portion is, from first to last, the front ear axle pin, the rear ear axle pin, and the side support rod axle pin.

[0017] In order to achieve the above technical effects, the present invention also provides an aircraft landing gear structure design system based on fatigue damage, comprising:

[0018] A data acquisition module, used to acquire a historical flight data set of the same type of aircraft, wherein the data set includes horizontal component velocity data and sinking component velocity data when the aircraft of the same type completes a landing;

[0019] A stress spectrum extraction module is used to construct a three-dimensional finite element analysis model of the model aircraft according to the size structure of the model aircraft, and use different horizontal component velocity data and different sinking component velocity data as combined boundary conditions, and use the three-dimensional finite element analysis model to analyze and obtain the stress spectrum of the landing gear separation part of the model aircraft during the process from taxiing takeoff to landing under each combined boundary condition; the separation part includes the front trunnion pin, the rear trunnion pin and the side support rod axle pin of the landing gear;

[0020] A model building module is used to statistically extract the stress cycle combination of maximum stress-0-maximum stress by using the rain flow counting method according to the maximum stress in each stress spectrum; fatigue life analysis is carried out on the three-dimensional finite element analysis model using each stress cycle combination to obtain the maximum shear strain and cycle number of the corresponding severance position under each stress cycle combination, and a strain fatigue analysis model related to the maximum shear strain and cycle number is constructed based on the Manson-Coffin formula;

[0021] A first analysis module is used to analyze and obtain a first limit load analysis value of each severance part of the aircraft to be designed according to the strain fatigue analysis model, the shear strain versus shear stress curve of the standard rod of the severance part material, and the number of design cycles in each severance part of the aircraft to be designed;

[0022] The second analysis module is used to obtain the maximum value of the horizontal component velocity within the range of ±3 standard deviations in the horizontal component velocity data distribution and the maximum value of the sinking component velocity within the range of ±3 standard deviations in the sinking component velocity data distribution by using a statistical method; extract the second limit load analysis value of each separation position of the landing gear with the maximum value of the horizontal component velocity and the maximum value of the sinking component velocity as the combined boundary conditions from the stress spectrum;

[0023] The design module is used to select the design shear load upper limit value of the corresponding separation part within the interval formed by the second limit load analysis value and the first limit load analysis value of the corresponding separation part according to the required separation order of each separation part, perform the separation part structure design, and complete the aircraft landing gear structure design.

[0024] Furthermore, in the model building module, the strain fatigue analysis model related to the maximum shear strain and the number of cycles constructed based on the Manson-Coffin formula is: ,in The maximum stress of the standard rod of the material at the separation site The shear strain amplitude under is the fatigue strength coefficient of the material at the detached part, is the shear modulus of the material at the detachment site, is the corresponding number of cycles, The fatigue ductility coefficient of the material corresponding to the separation part, , is the exponent obtained by fitting the data.

[0025] Furthermore, the first analysis module is used to substitute the design cycle number in the corresponding separation part of the aircraft to be designed into the strain fatigue analysis model, and jointly solve the shear strain versus shear stress curve of the standard rod material of the separation part to obtain the first limit load analysis value of the corresponding separation part.

[0026] Furthermore, in the design module, the required disconnection order of each disconnection part is, from first to last, the front ear axle pin, the rear ear axle pin, and the side support rod axle pin.

[0027] To achieve the above technical effects, the present invention also provides an application of an aircraft landing gear structure design method based on fatigue damage, and the aircraft landing gear structure design method based on fatigue damage is applied to obtain an aircraft landing gear structure.

[0028] Compared with the prior art, the present invention has the following beneficial effects:

[0029] 1. The present invention obtains a first limit load analysis value that takes into account fatigue damage of the detached part under alternating stress cycles by performing fatigue life analysis on the detached part; obtains a second limit load analysis value of the detached part based on historical flight data analysis of the same type of aircraft, the second limit load analysis value taking into account the combination conditions of the limit horizontal component velocity and the limit sinking component velocity under about 99% or more of normal landing conditions; selects the upper limit value of the design shear load of the detached part from the first limit load analysis value and the second load analysis value, while ensuring that the fatigue life of the detached part meets the design requirements, it also avoids the problem of a low upper limit value of the shear load of the detached part, and further ensures that the detached part of the aircraft to be designed is not prone to non-overload detachment during normal landing.

[0030] 2. Before constructing the strain fatigue analysis model, the stress spectrum extraction process of the landing gear separation part during the aircraft's taxiing takeoff to landing took into account the historical data of the horizontal velocity and sinking velocity of the same type of aircraft during landing, avoiding the traditional aircraft landing gear design that only considers the sinking speed while ignoring the impact effect of the aircraft's horizontal speed on the landing gear, ensuring the accuracy of the constructed strain fatigue analysis model. BRIEF DESCRIPTION OF THE DRAWINGS

[0031] Figure 1 A flow chart of the aircraft landing gear structure design method based on fatigue damage in Example 1 or 2;

[0032] Figure 2 This is a structural block diagram of the aircraft landing gear structure design system based on fatigue damage in Example 1;

[0033] Figure 3 A schematic diagram of the aircraft landing gear structure in Example 1 or 2;

[0034] Among them, 1. front ear axle pin; 2. rear ear axle pin; 3. side support rod axle pin; 4. data acquisition module; 5. stress spectrum extraction module; 6. model building module; 7. first analysis module; 8. second analysis module; 9. design module. DETAILED DESCRIPTION

[0035] The present invention is further described in detail below in conjunction with the embodiments and drawings. However, it should not be understood that the scope of the above subject matter of the present invention is limited to the following embodiments, and all technologies realized based on the content of the present invention belong to the scope of the present invention.

[0036] The present invention is further described below in conjunction with specific embodiments.

[0037] Example 1

[0038] See also Figure 1-Figure 3 , aircraft landing gear structure design method based on fatigue damage, including:

[0039] Acquire a historical flight data set of the same type of aircraft, wherein the data set includes horizontal component velocity data and sinking component velocity data of the aircraft during a complete landing;

[0040] According to the size structure of the model aircraft, a three-dimensional finite element analysis model of the model aircraft is constructed, and different horizontal component velocity data and different sinking component velocity data are used as combined boundary conditions, and the three-dimensional finite element analysis model is used to analyze and obtain the stress spectrum of the landing gear separation part of the model aircraft during the process from taxiing takeoff to landing under each combined boundary condition; the separation part includes the front trunnion pin 1, the rear trunnion pin 2 and the side support rod axle pin 3 of the landing gear;

[0041] According to the maximum stress in each stress spectrum, the stress cycle combination of maximum stress-0-maximum stress is statistically extracted by using the rain flow counting method; fatigue life analysis is carried out on the three-dimensional finite element analysis model using each stress cycle combination to obtain the maximum shear strain and cycle number of the corresponding severance position under each stress cycle combination, and a strain fatigue analysis model related to the maximum shear strain and cycle number is constructed based on the Manson-Coffin formula;

[0042] According to the strain fatigue analysis model, the shear strain versus shear stress curve of the standard rod of the material at the separation site, and the number of design cycles in each separation site of the aircraft to be designed, the first limit load analysis value of each separation site of the aircraft to be designed is obtained by analysis;

[0043] A statistical method is used to obtain the maximum value of the horizontal component velocity within the range of ±3 standard deviations in the horizontal component velocity data distribution, and the maximum value of the sinking component velocity within the range of ±3 standard deviations in the sinking component velocity data distribution; and a second limit load analysis value of each separation position of the landing gear with the maximum value of the horizontal component velocity and the maximum value of the sinking component velocity as combined boundary conditions is extracted from the stress spectrum;

[0044] According to the required separation order of each separation part, the design shear load upper limit value of the corresponding separation part is selected within the interval formed by the second limit load analysis value and the first limit load analysis value of the corresponding separation part, the separation part structure design is carried out, and the aircraft landing gear structure design is completed.

[0045] In this embodiment, by constructing a strain fatigue analysis model of the aircraft landing gear detachment site, the first limit load analysis value considering the fatigue damage of the detachment site under the alternating stress cycle can be obtained; and before the strain fatigue analysis model is constructed, the stress spectrum of the detachment site of the aircraft during the taxiing takeoff and landing process takes into account the historical data of the horizontal component velocity and the sinking component velocity of the same type of aircraft during landing, thereby avoiding the traditional aircraft landing gear design that only considers the sinking speed and ignores the impact effect of the horizontal speed of the aircraft on the landing gear, thereby ensuring the accuracy of the constructed strain fatigue analysis model. In addition, through statistical analysis, the second limit load analysis value corresponding to the combined boundary conditions of the maximum horizontal component velocity and the maximum sinking component velocity within the range of ±3 standard deviations is obtained. The second limit load analysis value takes into account the combination conditions of the limit horizontal component velocity and the limit sinking component velocity under about more than 99% of normal landing conditions. By selecting the upper limit value of the design shear load of the separation part from the first limit load analysis value and the second load analysis value, while ensuring that the fatigue life of the separation part meets the design requirements, the problem of a low upper limit value of the shear load of the separation part is avoided, further ensuring that the separation part of the aircraft to be designed is not prone to non-overload separation during normal landing.

[0046] Based on the same inventive concept, this embodiment also provides an aircraft landing gear structure design system based on fatigue damage, including:

[0047] Data acquisition module 4, used to acquire historical flight data set of the same model of aircraft, the data set including horizontal component velocity data and sinking component velocity data of the model of aircraft during complete landing;

[0048] The stress spectrum extraction module 5 is used to construct a three-dimensional finite element analysis model of the model aircraft according to the size structure of the model aircraft, and use the different horizontal component velocity data and the different sinking component velocity data as combined boundary conditions to analyze and obtain the stress spectrum of the landing gear separation part of the model aircraft from taxiing takeoff to landing under each combined boundary condition by using the three-dimensional finite element analysis model; the separation part includes the front trunnion pin 1, the rear trunnion pin 2 and the side support rod axle pin 3 of the landing gear;

[0049] Model construction module 6 is used to statistically extract the stress cycle combination of maximum stress-0-maximum stress by using the rain flow counting method according to the maximum stress in each stress spectrum; use each stress cycle combination to carry out fatigue life analysis on the three-dimensional finite element analysis model to obtain the maximum shear strain and cycle number of the corresponding severance position under each stress cycle combination, and construct a strain fatigue analysis model related to the maximum shear strain and cycle number based on the Manson-Coffin formula;

[0050] The first analysis module 7 is used to analyze and obtain the first limit load analysis value of each separation part of the aircraft to be designed according to the strain fatigue analysis model, the shear strain and shear stress variation curve of the standard rod of the separation part material, and the number of design cycles in each separation part of the aircraft to be designed;

[0051] The second analysis module 8 is used to obtain the maximum value of the horizontal component velocity within the range of ±3 standard deviations in the horizontal component velocity data distribution and the maximum value of the sinking component velocity within the range of ±3 standard deviations in the sinking component velocity data distribution by using a statistical method; extract the second limit load analysis value of each separation position of the landing gear with the maximum value of the horizontal component velocity and the maximum value of the sinking component velocity as the combined boundary conditions from the stress spectrum;

[0052] Design module 9 is used to select the design shear load upper limit value of the corresponding separation part within the interval formed by the second limit load analysis value and the first limit load analysis value of the corresponding separation part according to the required separation order of each separation part, perform the separation part structure design, and complete the aircraft landing gear structure design.

[0053] Example 2

[0054] See also Figure 1 and Figure 3 This embodiment takes the design of a certain type of aircraft landing gear structure to be designed as an example to explain in detail the process of the aircraft landing gear structure design method based on fatigue damage of the present invention. The specific design steps of the aircraft to be designed are as follows:

[0055] Step 1, obtaining a historical flight data set of the same type of aircraft, wherein the data set includes horizontal component velocity data and sinking component velocity data of the aircraft when the aircraft has completely landed;

[0056] Step 2: construct a three-dimensional finite element analysis model of the model aircraft according to the size structure of the model aircraft, and use different horizontal component velocity data and different sinking component velocity data as combined boundary conditions, and use the three-dimensional finite element analysis model to analyze and obtain the stress spectrum of the landing gear separation part of the model aircraft from taxiing takeoff to landing under each combined boundary condition; the separation part includes the front trunnion pin 1, the rear trunnion pin 2 and the side support rod axle pin 3 of the landing gear;

[0057] Step 3: According to the maximum stress in each stress spectrum, a stress cycle combination of maximum stress-0-maximum stress is statistically extracted by using a rain flow counting method; fatigue life analysis is carried out on the three-dimensional finite element analysis model using each stress cycle combination to obtain the maximum shear strain and number of cycles of the corresponding severance position under each stress cycle combination, and a strain fatigue analysis model related to the maximum shear strain and the number of cycles is constructed based on the Manson-Coffin formula;

[0058] In this embodiment, the strain fatigue analysis model related to the maximum shear strain and the number of cycles constructed based on the Manson-Coffin formula is: ,in The maximum stress of the standard rod of the material at the separation site The shear strain amplitude under is the fatigue strength coefficient of the material at the detached part, is the shear modulus of the material at the detachment site, is the corresponding number of cycles, The fatigue ductility coefficient of the material corresponding to the separation part, , is the exponent obtained by fitting the data.

[0059] Step 4: Analyze and obtain the first limit load analysis value of each severance part of the aircraft to be designed according to the strain fatigue analysis model, the shear strain versus shear stress curve of the standard rod of the severance part material, and the number of design cycles in each severance part of the aircraft to be designed;

[0060] In this embodiment, the design cycle number in the corresponding separation part of the aircraft to be designed is substituted into the strain fatigue analysis model, and the first limit load analysis value of the corresponding separation part is obtained by jointly solving the shear strain versus shear stress curve of the standard rod material of the separation part.

[0061] Step 5: using a statistical method to obtain the maximum value of the horizontal component velocity within the range of ±3 standard deviations in the horizontal component velocity data distribution, and the maximum value of the sinking component velocity within the range of ±3 standard deviations in the sinking component velocity data distribution; extracting the second limit load analysis value of each separation position of the landing gear with the maximum value of the horizontal component velocity and the maximum value of the sinking component velocity as the combined boundary conditions from the stress spectrum;

[0062] Step 6: According to the required separation order of each separation part, select the design shear load upper limit value of the corresponding separation part within the interval formed by the second limit load analysis value and the first limit load analysis value of the corresponding separation part, perform the separation part structure design, and complete the aircraft landing gear structure design;

[0063] In this embodiment, the required disconnection order of each disconnection part is, from first to last, the front ear axle pin 1, the rear ear axle pin 2, and the side support rod axle pin 3; for example, if the value range of the design shear load upper limit value of the front ear axle pin 1 is calculated to be [a, b], the value range of the design shear load upper limit value of the rear ear axle pin 2 is [c, d], and the value range of the design shear load upper limit value of the side support rod axle pin 3 is [e, f], when designing the disconnection order, a value a+δ1 can be selected from [a, b] as the design shear load upper limit value of the front ear axle pin 1, a value c+δ1+δ2 can be selected from [c, d] as the design shear load upper limit value of the rear ear axle pin 2, and a value e+δ1+δ2+δ3 can be selected from [e, f] as the design shear load upper limit value of the side support rod axle pin 3, where δ1<δ2<δ3, to ensure that the disconnection order meets the design order requirements.

[0064] The above-mentioned specific implementation can be partially adjusted in different ways by those skilled in the art without departing from the principle and purpose of the present invention. The protection scope of the present invention shall be based on the claims and shall not be limited by the above-mentioned specific implementation. Each implementation scheme within its scope shall be subject to the constraints of the present invention.

Claims

1. A fatigue damage-based aircraft landing gear structure design method, characterized in that: include: Acquire a historical flight data set of the same type of aircraft, wherein the data set includes horizontal component velocity data and sinking component velocity data of the aircraft during a complete landing; According to the size structure of the model aircraft, a three-dimensional finite element analysis model of the model aircraft is constructed, and different horizontal component velocity data and different sinking component velocity data are used as combined boundary conditions, and the three-dimensional finite element analysis model is used to analyze and obtain the stress spectrum of the landing gear separation part of the model aircraft during the process from taxiing takeoff to landing under each combined boundary condition; the separation part includes the front trunnion pin, the rear trunnion pin and the side support rod axle pin of the landing gear; According to the maximum stress in each stress spectrum, a stress cycle combination of maximum stress-0-maximum stress is statistically extracted by using a rain flow counting method; Using each stress cycle combination, fatigue life analysis is carried out on the three-dimensional finite element analysis model to obtain the maximum shear strain and cycle number of the corresponding severance position under each stress cycle combination, and a strain fatigue analysis model related to the maximum shear strain and cycle number is constructed based on the Manson-Coffin formula; According to the strain fatigue analysis model, the shear strain versus shear stress curve of the standard rod of the material at the separation site, and the number of design cycles in each separation site of the aircraft to be designed, the first limit load analysis value of each separation site of the aircraft to be designed is obtained by analysis; A statistical method is used to obtain the maximum value of the horizontal component velocity within the range of ±3 standard deviations in the horizontal component velocity data distribution, and the maximum value of the sinking component velocity within the range of ±3 standard deviations in the sinking component velocity data distribution; and a second limit load analysis value of each separation position of the landing gear with the maximum value of the horizontal component velocity and the maximum value of the sinking component velocity as combined boundary conditions is extracted from the stress spectrum; According to the required separation order of each separation part, the design shear load upper limit value of the corresponding separation part is selected within the interval formed by the second limit load analysis value and the first limit load analysis value of the corresponding separation part, the separation part structure design is carried out, and the aircraft landing gear structure design is completed.

2. The method for designing aircraft landing gear structure based on fatigue damage according to claim 1, characterized in that: The strain fatigue analysis model related to the maximum shear strain and the number of cycles constructed based on the Manson-Coffin formula is: ,in The maximum stress of the standard rod of the material at the separation site The shear strain amplitude under is the fatigue strength coefficient of the material at the detached part, is the shear modulus of the material at the detachment site, is the corresponding number of cycles, The fatigue ductility coefficient of the material corresponding to the separation part, , is the exponent obtained by fitting the data.

3. The method for designing aircraft landing gear structure based on fatigue damage according to claim 1, characterized in that: The method for obtaining the first limit load analysis value of each breakaway portion of the aircraft to be designed includes: The design cycle number in the corresponding severance part of the aircraft to be designed is substituted into the strain fatigue analysis model, and the first limit load analysis value of the corresponding severance part is obtained by jointly solving the shear strain versus shear stress curve of the standard bar material of the severance part.

4. The method for designing aircraft landing gear structure based on fatigue damage according to claim 1, characterized in that: The required disconnection order of each disconnection part is, from first to last, the front ear axle pin, the rear ear axle pin, and the side support rod axle pin.

5. The aircraft landing gear structure design system based on fatigue damage is characterized by: include: A data acquisition module, used to acquire a historical flight data set of the same type of aircraft, wherein the data set includes horizontal component velocity data and sinking component velocity data when the aircraft of the same type completes a landing; A stress spectrum extraction module is used to construct a three-dimensional finite element analysis model of the model aircraft according to the size structure of the model aircraft, and use different horizontal component velocity data and different sinking component velocity data as combined boundary conditions, and use the three-dimensional finite element analysis model to analyze and obtain the stress spectrum of the landing gear separation part of the model aircraft during the process from taxiing takeoff to landing under each combined boundary condition; the separation part includes the front trunnion pin, the rear trunnion pin and the side support rod axle pin of the landing gear; A model building module is used to statistically extract a stress cycle combination of maximum stress-0-maximum stress by using a rain flow counting method according to the maximum stress in each stress spectrum; Using each stress cycle combination, fatigue life analysis is carried out on the three-dimensional finite element analysis model to obtain the maximum shear strain and cycle number of the corresponding severance position under each stress cycle combination, and a strain fatigue analysis model related to the maximum shear strain and cycle number is constructed based on the Manson-Coffin formula; A first analysis module is used to analyze and obtain a first limit load analysis value of each severance part of the aircraft to be designed according to the strain fatigue analysis model, the shear strain versus shear stress curve of the standard rod of the severance part material, and the number of design cycles in each severance part of the aircraft to be designed; The second analysis module is used to obtain the maximum value of the horizontal component velocity within the range of ±3 standard deviations in the horizontal component velocity data distribution and the maximum value of the sinking component velocity within the range of ±3 standard deviations in the sinking component velocity data distribution by using a statistical method; extract the second limit load analysis value of each separation position of the landing gear with the maximum value of the horizontal component velocity and the maximum value of the sinking component velocity as the combined boundary conditions from the stress spectrum; The design module is used to select the design shear load upper limit value of the corresponding separation part within the interval formed by the second limit load analysis value and the first limit load analysis value of the corresponding separation part according to the required separation order of each separation part, perform the separation part structure design, and complete the aircraft landing gear structure design.

6. The aircraft landing gear structure design system based on fatigue damage according to claim 5, characterized in that: In the model building module, the strain fatigue analysis model related to the maximum shear strain and the number of cycles constructed based on the Manson-Coffin formula is: ,in The maximum stress of the standard rod of the material at the separation site The shear strain amplitude under is the fatigue strength coefficient of the material at the detached part, is the shear modulus of the material at the detachment site, is the corresponding number of cycles, The fatigue ductility coefficient of the material corresponding to the separation part, , is the exponent obtained by fitting the data.

7. The aircraft landing gear structure design system based on fatigue damage according to claim 5, characterized in that: The first analysis module is used to substitute the design cycle number in the corresponding separation part of the aircraft to be designed into the strain fatigue analysis model, and jointly solve the shear strain versus shear stress variation curve of the standard rod material of the separation part to obtain the first limit load analysis value of the corresponding separation part.

8. The aircraft landing gear structure design system based on fatigue damage according to claim 5, characterized in that: In the design module, the required disconnection order of each disconnection part is, from first to last, the front ear axle pin, the rear ear axle pin, and the side support rod axle pin.

9. An application of a fatigue damage-based aircraft landing gear structure design method, characterized in that: Apply the aircraft landing gear structure design method based on fatigue damage as described in any one of claims 1 to 4 to obtain an aircraft landing gear structure.

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