A rapid assessment method for vibration fatigue life of aircraft structures

Abstract

The present invention discloses a method for rapid assessment of aircraft structure vibration fatigue life. Based on the fatigue DFR calculation method and flight test measured data, a new vibration fatigue equal amplitude load / stress is quickly obtained by using a load / stress superposition method, thereby obtaining the number of fatigue cycles and the duration of the vibration cycle under the corresponding vibration environment. The present invention can effectively shorten the assessment cycle and ensure flight test safety.

Description

Technical Field

[0001] The invention belongs to the field of aircraft structure fatigue strength calculation, and in particular relates to a method for quickly evaluating the vibration fatigue life of an aircraft structure. Background Art

[0002] Newly developed aircraft generally have various vibration problems during the test flight. On the one hand, it may be caused by unreasonable design factors such as insufficient stiffness. On the other hand, it may be caused by inaccurate load estimation. Generally, it occurs in the rudder structure. For these vibrations, an evaluation opinion needs to be given. On the one hand, it is necessary to ensure the safety of the aircraft and on the other hand, it is necessary to ensure the normal progress of the test flight. Traditional vibration analysis methods are generally based on measured data and are divided into time domain method and frequency domain method. The time domain method needs to be based on a large amount of processing of measured data. The frequency domain method is based on power spectrum density and needs to establish a finite element model, perform frequency response analysis, and set fatigue analysis parameters. The process is relatively complicated and the calculation cycle is long. Summary of the invention

[0003] The present invention aims at solving the above problems and proposes a method for quickly evaluating the vibration fatigue life of aircraft structures. The present invention can effectively shorten the evaluation period and ensure the safety of flight tests.

[0004] Technical solution: A method for rapid assessment of aircraft structure vibration fatigue life, based on fatigue DFR calculation method and flight test data, uses load / stress superposition to quickly obtain new vibration fatigue equal-amplitude load / stress, thereby obtaining the number of fatigue cycles and vibration cycle duration under the corresponding vibration environment.

[0005] The aforementioned rapid assessment method for aircraft structure vibration fatigue life has the following steps:

[0006] (1) Obtaining the measured time domain data of the vibration part, determining the vibration overload based on the time domain data, and obtaining the vibration frequency of the vibration part based on the time domain-frequency domain conversion;

[0007] (2) Analyze the detailed fatigue ratings and corresponding load / stress spectra corresponding to the vibration locations;

[0008] (3) Calculate the inertial load / stress caused by vibration at the vibrating part, i.e., F v / S v ;

[0009] (4) superimposing the maximum fatigue load / stress of the vibrating part when it is not vibrating and the inertia load / stress in step (3) to form a new fatigue constant amplitude load / stress;

[0010] (5) calculating fatigue life based on the detailed fatigue rating in step (2) and the fatigue constant amplitude load / stress in step (2);

[0011] (6) According to the fatigue life in step (5) and the vibration frequency in step (1), the vibration cycle duration under vibration overload in step (1) is obtained.

[0012] In the aforementioned rapid assessment method for the vibration fatigue life of aircraft structures, in step (1), for random loads, the vibration cycle durations corresponding to several typical vibration overload values ​​in the random load are first obtained according to steps (1) to (6), and then the vibration cycle duration under the random vibration is obtained by interpolation.

[0013] In the aforementioned method for rapid assessment of aircraft structure vibration fatigue life, in step (1), the measured time domain data is data collected by an overload sensor at the vibration location, and the collection principle is as follows: the sensor collects data in the direction of large vibration overload.

[0014] In the aforementioned rapid assessment method for aircraft structure vibration fatigue life, in step (2), the detailed fatigue rating is calculated according to the corresponding calculation method based on the structure type, and the structure type includes openings, ears, and connectors; the load / stress spectrum is used to screen out the maximum fatigue load / stress, that is, F max / S max When there is no fatigue load / stress spectrum data, the limit load of 0.8 times the static strength calculation is conservatively used as the maximum fatigue load / stress.

[0015] In the aforementioned rapid assessment method for the vibration fatigue life of aircraft structures, in step (3), when calculating the inertial load / stress, the load is determined by multiplying the structure weight and the vibration overload, and the stress is converted into the corresponding overload when calculating the overload and static strength.

[0016] In the aforementioned rapid assessment method for vibration fatigue life of aircraft structures, in step (5), the fatigue life is calculated based on the detailed fatigue rating and fatigue constant amplitude load / stress, and considering the fatigue reliability factor FRF=1.5.

[0017] In the aforementioned method for rapid assessment of aircraft structure vibration fatigue life, in step (6), the duration of the vibration cycle under the overload and frequency is calculated based on the calculated fatigue life and vibration frequency.

[0018] Beneficial effects: Based on the fatigue strength DFR calculation method, the present invention designs a rapid vibration fatigue assessment method, which is mainly based on the fatigue life assessment and measured data of the aircraft structure, and adopts the load / stress superposition method to give the vibration period of the aircraft structure, which is used for flight test safety assessment and determination of aircraft structure vibration limits.

[0019] The present invention is based on the factors related to aircraft structure fatigue analysis: DFR and fatigue load spectrum; based on the measured data of aircraft structure vibration, the inertial load caused by vibration and the maximum load in the original fatigue load spectrum are superimposed to form a new equal-amplitude fatigue load, and the fatigue vibration cycle under the vibration environment is obtained according to the new equal-amplitude stress and structural DFR value; the evaluation method of the present invention is efficient and easy to implement, which can ensure the safety of the aircraft and the normal test flight, and is applicable to all newly developed aircraft; in addition, the present invention can also be used for the preliminary determination of the vibration limit of the aircraft structure. It has been verified that the present invention can improve the efficiency of vibration fatigue analysis by about 50% while ensuring safety.

[0020] The present invention provides a method for quickly evaluating the vibration fatigue life of an aircraft structure, which overcomes the disadvantages of the traditional vibration fatigue calculation process being complicated, cumbersome, and having many input parameters, and superimposes the inertial load caused by vibration into the traditional fatigue analysis. The method is simple, fast, and has a clear process. The method can be used to guide the evaluation of the fatigue life of a structure when vibration occurs during a flight test, and can effectively shorten the evaluation cycle and ensure the safety of the flight test; the method can be used to guide the determination of the vibration limits of various parts of an aircraft, allowing a certain amount of vibration to exist while meeting the requirements of the fatigue life of the aircraft.

[0021] The following examples are taken as examples to further illustrate the beneficial effects of the present invention:

[0022] The vertical overload of the aileron structure vibration is 8.8g, the vibration frequency is 98Hz, the weight of the aileron is 70kg, the aileron has 2 suspension joints DFR=89.6, 0.8 times the limit load is the heading load Fx=15666N, the vertical load Fz=5738N, the vertical load caused by the vibration of a single joint is Fz=3080N, after superimposing the load, Fx=15666N, Fz=8818N and Fx=15666N, Fz=2658N, the net area of ​​the ear is 200mm 2 After superposition, the stress is Smax=87.4MPa, Smin=77.3MPa. The fatigue life cycle number under the new load spectrum is calculated to be 47667122. At this frequency and overload, the corresponding continuous vibration failure time is 135 hours. BRIEF DESCRIPTION OF THE DRAWINGS

[0023] Figure 1 The schematic diagram of the aileron structure sensor in the embodiment;

[0024] Figure 2 It is a schematic diagram of the aircraft structure vibration time domain measured data and frequency domain data in the embodiment;

[0025] Figure 3 is a schematic diagram of frequency domain data of aircraft structure vibration in an embodiment;

[0026] Figure 4Schematic diagram of the vibration fatigue detail structure in the embodiment. DETAILED DESCRIPTION

[0027] In order to make the implementation process of the present invention clearer, the technical scheme in the embodiment of the present invention will be described in more detail below in conjunction with the drawings in the embodiment of the present invention. In the drawings, the same or similar reference numerals throughout represent the same or similar elements or elements with the same or similar functions. The described embodiments are part of the embodiments of the present invention, not all of the embodiments, and the undetailed parts are all conventional technologies. The embodiments described below with reference to the drawings are exemplary and are intended to be used to explain the present invention, and should not be construed as limitations on the present invention. Based on the embodiments in the present invention, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present invention.

[0028] Example 1. Figure 1 As shown, it is a schematic diagram of time domain vibration involved in the rapid assessment method of aircraft structure vibration fatigue life of the present invention.

[0029] The method for rapid assessment of mechanical structure vibration fatigue life of the present invention is specifically implemented as follows:

[0030] Step 1: Obtain the measured time domain data of aircraft structure vibration, see Figure 2 ;

[0031] Step 2: Convert the measured time domain data into frequency domain to obtain the vibration frequency, see Figure 2 ;

[0032] Step 3: Analyze the structural fatigue detail DFR (detailed fatigue rating), see Figure 3 ;

[0033] Step 4: Obtain the fatigue details of the structure under normal fatigue conditions. Maximum fatigue load / stress F max / S max ;

[0034] Step 5: Calculate the load / stress (F 2 max / S 2 max);

[0035] Step 6: Superimpose normal fatigue load / stress (F max / S max ) and loads / stresses due to inertia (±F v / S v ), forming a new load / stress of equal amplitude;

[0036] Step 7: Calculate the number of fatigue life cycles based on the new constant amplitude stress and fatigue detail DFR value;

[0037] Step 8: Calculate the duration of the vibration cycle according to the frequency corresponding to the overload and the number of fatigue life cycles;

[0038] Step 9. Calculate the typical overload for random vibration loads, repeat the above steps, and interpolate to determine the period length under random vibration based on the proportion of typical overload occurrence.

[0039] The above is only a specific embodiment of the present invention, but the protection scope of the present invention is not limited thereto. Any changes or substitutions that can be easily thought of by a person skilled in the art within the technical scope disclosed by the present invention should be included in the protection scope of the present invention. Therefore, the protection scope of the present invention shall be based on the protection scope of the claims.

Claims

1. A method for rapid assessment of aircraft structure vibration fatigue life, characterized in that: Based on the fatigue DFR calculation method and flight test data, the new vibration fatigue equal-amplitude load / stress is quickly obtained by using the load / stress superposition method, so as to obtain the fatigue cycle number and vibration cycle duration under the corresponding vibration environment; the steps are as follows: (1) Obtaining the measured time domain data of the vibration part, determining the vibration overload based on the time domain data, and obtaining the vibration frequency of the vibration part based on the time domain-frequency domain conversion; (2) Analyze the detailed fatigue ratings and corresponding load / stress spectra corresponding to the vibration locations; (3) Calculate the inertial load / stress caused by vibration at the vibrating part, i.e., F v / S v ; (4) superimposing the maximum fatigue load / stress of the vibrating part when it is not vibrating and the inertia load / stress in step (3) to form a new fatigue constant amplitude load / stress; (5) calculating fatigue life based on the detailed fatigue rating in step (2) and the fatigue constant amplitude load / stress in step (4); (6) According to the fatigue life in step (5) and the vibration frequency in step (1), the vibration cycle duration under vibration overload in step (1) is obtained.

2. The method for rapid assessment of aircraft structure vibration fatigue life according to claim 1, characterized in that: In step (1), the measured time domain data is data collected by the overload sensor at the vibration location, and the collection principle is as follows: the sensor collects data in the direction of the large vibration overload.

3. The method for rapid assessment of aircraft structure vibration fatigue life according to claim 1, characterized in that: In step (2), the detailed fatigue rating is calculated according to the structure type and the corresponding calculation method, and the structure type includes openings, ears, and connectors; Load / stress spectrum, used to screen out the maximum fatigue load / stress, i.e. F max / S max When there is no fatigue load / stress spectrum data, the limit load of 0.8 times the static strength calculation is conservatively used as the maximum fatigue load / stress.

4. The method for rapid assessment of aircraft structure vibration fatigue life according to claim 1, characterized in that: In step (3), when calculating the inertial load / stress, the load is determined by multiplying the structure weight and the vibration overload, and the stress is converted into stress according to the corresponding overload when calculating the overload and static strength.

5. The method for rapid assessment of aircraft structure vibration fatigue life according to claim 1, characterized in that: In step (5), the fatigue life is calculated based on the detailed fatigue rating and fatigue constant amplitude load / stress, and considering the fatigue reliability factor FRF=1.

5.

6. The method for rapid assessment of aircraft structure vibration fatigue life according to claim 1, characterized in that: In step (6), the duration of the vibration period at the overload and frequency is calculated based on the calculated fatigue life and vibration frequency.

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