KR20130118065A - Evaluation methods for s-n fatigue property combined with damages induced by aging effect - Google Patents

Abstract

PURPOSE: An S-N fatigue property evaluating method according to a shape of damage to the surface of a component of a long-term operation aircraft is provided to metallically analyze various kinds of corrosive damage to primary structures of a superannuated aircraft and to measure the fatigue lifetime of actual components, thereby predicting fatigue lifetime according to a shape of damage to the components of the aircraft. CONSTITUTION: An S-N fatigue property evaluating method according to a shape of damage to the surface of a component of a long-term operation aircraft includes the following steps of: selecting target components of the long-term operation aircraft for predicting the remaining fatigue lifetime of structures of the aircraft; building a database in respect to a shape of the corrosion of the target components separated from the long-term operation aircraft; sampling an S-N fatigue testing piece from the target components; performing a fatigue test by using the S-N fatigue testing piece; observing a portion on which initial fatigue cracks are generated by using a fatigue fractured surface after the S-N fatigue test; and predicting the fatigue lifetime by using a crack growth analysis code.

Description

Evaluation Method of S-N Fatigue Characteristics According to Surface Damage Patterns of Aircraft Components for Long-Term Operation {EVALUATION METHODS FOR S-N FATIGUE PROPERTY COMBINED WITH DAMAGES INDUCED BY AGING EFFECT}

The present invention relates to a method for evaluating SN fatigue characteristics according to the surface damage pattern of a long-term operation aircraft, and more specifically, in order to more accurately manage the life of the aircraft airframe, various corrosion damages generated in an aging aircraft are applied. The present invention relates to a method for evaluating SN fatigue characteristics according to surface damage patterns of long-term operation aircraft parts by analyzing the fatigue life of real parts and evaluating SN fatigue characteristics according to damage forms.

Currently, some aircraft that are in use beyond their initial design service life are inevitably required for additional 20 years or more, depending on their economic and military operational needs. These aircraft cause various forms of corrosion, wear, and fatigue damage according to the accumulation of flight time, which is a factor that lowers the stability and operation rate of the aircraft. In the case of corrosion in particular, no specific criteria are given during the inspection cycle, and the proven methods for the lifetime and risk of corrosion structures are quite limited.

Accordingly, the present invention provides a fatigue analysis according to the form of damage to aircraft parts by measuring the fatigue life of the seal parts in a metallographic analysis of various corrosion damages occurring in the main structures of the aging aircraft for more accurate life management of long-term operation aircraft. It is an object to provide a method for predicting lifespan.

In order to achieve the above object, the SN fatigue characteristic evaluation method according to the long-term operating aircraft component surface damage pattern according to the present invention, in order to predict the residual fatigue life of the structure of the long-term operating aircraft, Selecting step (S100), and the database of the surface corrosion pattern of the target component detached from the long-term operation aircraft (S200), and processing the SN fatigue test piece in the target component (S300), Performing a fatigue test using the SN fatigue test piece (S400), and after the SN fatigue test, using the fatigue fracture surface to observe the initial fatigue crack generation site (S500), and using the crack propagation analysis code It is characterized in that it comprises a step (S600) for predicting the fatigue life.

In addition, the SN fatigue characteristics evaluation method according to the long-term operation aircraft parts surface damage form according to the present invention, in the step (S100), the target parts are parts replaced by reaching the design life without cracking during operation of the aircraft It is characterized by that.

In addition, the SN fatigue characteristic evaluation method according to the long-term operating aircraft component surface damage form according to the present invention, in the step (S300), the SN fatigue test piece is characterized in that the plate-shaped fatigue specimen including a hole taken from the target component do.

In addition, the SN fatigue characteristic evaluation method according to the long-term operating aircraft component surface damage form according to the present invention, in the step (S400), the fatigue test, the stress ratio (minimum load / maximum load): 0.1, cycle: 40 Air, lab air, N _{run out} = 10 ⁷ cycles (run-out refers to the life of the specimen that does not break under any constant amplitude load).

In addition, the SN fatigue characteristic evaluation method according to the long-term operating aircraft component surface damage pattern according to the present invention, in the step (S600), using the crack propagation equation to predict the fatigue life, the crack propagation equation is the lower limit stress Applying NASGRO equation considering the expansion coefficient range and fracture toughness, the NASGRO equation is

Where N is the number of working fatigue loads, a is the crack length, R is the stress ratio, K is the stress intensity factor, C, n, p and q are the material constants, and f is the crack open function opening function).

In addition, the SN fatigue characteristic evaluation method according to the long-term operating aircraft component surface damage pattern according to the present invention, in the step (S500), the surface damage generated in the fatigue fracture surface affects the remaining fatigue life of the target component In order to judge, it is characterized by deriving EIFS (Equivalent Initial Flaw Size) value, which is not a crack but exhibits a crack growth behavior similar to a crack.

According to the present invention, for more accurate life management of long-term operation aircraft, by analyzing the metallurgical corrosion damage in the main structure of the aging aircraft and measuring the fatigue life of the seal parts, fatigue according to the type of aircraft parts damage This has the effect of providing a way to predict lifespan.

1 shows the SN fatigue characteristics of a 7075-T7351 specimen.
2 shows SN fatigue properties of a 15% spar structure.
3 is a diagram showing the relationship between the defect area measured at the fracture surface and the equivalent initial crack (EIFS).
4 shows the association between surface damage size and EIFS.
5 is a view showing the results of the crack growth analysis by deriving EIFS in 15% Wing Spar damaged specimens and surface polished specimens, respectively.
6 shows stress values for evaluating EIFS.

Hereinafter, preferred embodiments of the present invention will be described in detail with reference to the accompanying drawings and the following description. However, the present invention is not limited to the embodiments described herein but may be embodied in other forms. Rather, the embodiments disclosed herein are provided so that the disclosure can be thorough and complete, and will fully convey the scope of the invention to those skilled in the art. Like reference numerals designate like elements throughout the specification. It is to be understood that the terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the invention. In the present specification, the singular form includes plural forms unless otherwise specified in the specification. &Quot; comprises "and / or" comprising ", as used herein, unless the recited element, step, operation, and / Or additions.

In the present invention, for ease of explanation, by analyzing the metallurgical corrosion damage occurring in the main structure of the F-5 aircraft operating for more than 30 years and measuring the fatigue life of the seal parts, fatigue life according to the aircraft parts damage type Although the present invention is not limited thereto, the present invention can be used in a method for predicting fatigue life for all other aged aircraft.

First of all, in the present invention, digital data is collected by using optical instruments for surface corrosion, defects or damages in major parts detached from an actual aircraft, and quantification of corrosion forms is performed based on the measurement results. Representatively, the plate fatigue specimens centered on the holes were taken from 15% spar structure on the wing of the aircraft, and the SN fatigue test was performed, and the fracture surface was observed by scanning electron microscope (SEM). To quantify the shape of the corrosion damage. Based on the above data, the residual fatigue life is predicted using the flight load spectrum and the commercial fatigue analysis S / W, and compared with the experimental data.

SN fatigue characteristics evaluation method according to the long-term operating aircraft component surface damage pattern according to the present invention, in order to predict the residual fatigue life for the structure of the long-term operating aircraft, selecting the target component of the long-term operating aircraft (S100), Database (S200) a surface corrosion pattern of the target component detached from the long-term operating aircraft, processing a SN fatigue test specimen from the target component (S300), and fatigue using the processed SN fatigue test specimen. Performing the test (S400), after the SN fatigue test, the step of observing the initial fatigue crack generation site using the fatigue fracture surface (S500) and the step of predicting the fatigue life using the crack propagation analysis code (S600) ).

Experimental Method

1. Target parts

The present invention utilizes key components of an F-5E aircraft operated for more than 30 years to predict residual fatigue life for long operating aircraft structures. Therefore, the parts used are designed with the concept of safe-life, which means that no cracks occur during operation, but the parts that have been replaced after reaching the design life are 15% spar. Table 1 summarizes the part names, shapes, materials and flight times of the parts.

2. Material Properties and Chemical Composition

The aircraft structure is mainly used aluminum alloy material of 2xxx, 7xxx series, most of the F-5E aircraft components used in the present invention is Al 7xxx series mainly. The material of the parts used in the present invention is 7075-T7351 aluminum alloy, and the mechanical properties and chemical compositions are shown in Tables 2 and 3, respectively. Al 7075 alloy is a Cu-added alloy to increase the strength in Al-Zn-Mg alloys, which has the highest strength among aluminum alloys, but is easily prevented due to stress corrosion cracking. Measures are important. Therefore, by performing T73, T76 solution heat-treated and over aged on the Al 7075 alloy, even if the strength is reduced by about 10 to 20% compared to the case of T6 heat treatment, the stress corrosion cracking resistance is improved.

3. Test and analysis device

In the present invention, by correlating the surface corrosion patterns of the main parts removed from the actual aircraft, and performing a fatigue test using test specimens taken from the actual parts, the correlation between the corrosion damage occurring in the aircraft structure and the remaining fatigue life is identified. I would like to. Therefore, the equipment necessary to evaluate this is as follows.

1) Fatigue Tester: Instron's maximum capacity 10ton, 8801 series hydraulic fatigue tester is used, and the software is controlled by computer through the same console program.

2) Optical Microscope: LEICA's DMLM microscope allows observation and imaging up to 15 to 500 times, and more precise microstructure analysis through computer image analysis software connected to the microscope.

3) Real microscope: OLYMPUS's SZ61 microscope, which allows you to observe and photograph the desired position regardless of the specimen's position with a mobile microscope, and to observe the magnification of 200 times on the monitor, making it easier to find corrosion or cracks. It also has a fluid light source, so it is easy to shoot around the hole or rivet using the light source.

4) Scanning electron microscope: This is a scanning electron microscope of Jeol. It is mainly used for initial fatigue crack observation, and it can be observed up to about 18 to 300,000 times, and it is equipped with EDX analysis device. It is also possible.

4. SN under constant amplitude load Fatigue test

1) S-N fatigue test piece processing condition

In the 15% spar parts that reached the safe-life design life used in the present invention, plate-shaped fatigue specimens were taken around the hole, and the fatigue specimen shapes according to the components are shown in Table 4. Since plate-like specimens including holes were taken from the actual parts, the specimens had a dimensional difference little by little. In order to evaluate the residual fatigue life according to the long-term operating environment, the fatigue test was carried out in the as-cut state maintaining the existing surface state, and to compare the fatigue life with the existing alloy (bare), some specimens were surface polished Fatigue tests were performed in an as-polished state.

2) S-N Fatigue Test

Fatigue test was performed to evaluate the fatigue characteristics of the specimens with holes using the INSTRON 8801 equipment, detailed fatigue test conditions are as follows.

Stress ratio: 0.1 (Stress ratio is minimum load / maximum load)

Cycle: 40㎐

Lab air

N _{run out} = 10 ⁷ cycles (run-out refers to the life of a specimen that does not break under any constant amplitude load)

References: ASTM E466, MIL-HDBK-5J

Figure 1 shows the SN fatigue characteristics of the 7075-T7351 specimen without notches or holes, and in the case of as-polished without corrosion, it can be seen that the fatigue is about 200 MPa. In general, _Kt = 3.0 for the presence of holes, and fatigue life will be reduced by about three times. Thus, with reference to the ASTM and US military specifications, the added stress of the present invention was determined.

5. Observation of fracture surface

After the SN fatigue test, SEM was used to observe the site of the initial fatigue crack generation at the fatigue fracture surface, and the preparation of the specimen was cut to a thickness of about 10 mm at the fracture surface in a direction perpendicular to the load applied from the fractured specimen. After washing for 3 to 5 minutes in an ultrasonic cleaner containing alcohol, the specimen was fixed with a conductive tape to observe the fracture surface. In order to examine the fracture behavior of the entire specimen, the fracture surface observation is performed by sequentially tilting, mosaicing and attaching each of the low magnification (x15 ~ x25) photographs, and through high magnification (x500 ~ x2000) We wanted to know where the initial fatigue crack started and where it progressed.

6. Fatigue Life Prediction Using Crack Growth Analysis Code

In the present invention, commercial fatigue crack growth analysis S / W NASGRO ver. The fatigue analysis was performed using 6.01. The crack propagation equation is applied to NASGRO equation below considering the lower limit stress intensity range and fracture toughness.

Where N is the number of working fatigue loads, a is the crack length, R is the stress ratio, K is the stress intensity factor, and C, n, p and q are the material constants obtained empirically. In addition, f is a crack opening function, and the modified generalized willenborg equation proposed by Willenbporg is used to consider the effect of crack retardation due to overload in the analysis.

Findings and Discussion

1. Surface observation

Stereoscopic observations of surface periphery defects are made to verify how various corrosion damages appearing on the subject part are actually related to the residual fatigue life of the aircraft structure. Most of the observed surface corrosion such as pitting is mainly observed in the periphery including the holes, and their size was measured using a software called an image tool. Surface damage around the hole was observed in a magnified x20 size through a real microscope, and the cross-sectional area was measured and summarized as shown in Table 5 below. The size of the surface damage observed with a real microscope was about 0.05 to 0.20 mm2.

2. Fatigue Test Results

For the S-N fatigue test, the additional load was determined by referring to ASTM and the US Military Specification (MIL-HDBK-5J). The maximum added stress was 200 MPa and the stress ratio was 0.1. Fatigue test results for the wing spar are shown in Table 6 and Figure 2, where the added as-exposed specimens were lower in fatigue resistance compared to the as-polised specimens that were subjected to surface grinding after collection. It can be seen that it is much lower in the load range. When the fatigue limit is compared through the present invention, it can be seen that the fatigue resistance is reduced by an average of about 30% due to corrosion damage. After the fatigue test, low and high magnification observations are made to observe the initial fatigue crack generation for each specimen. As-exposed specimens from the wing spar were all cracked at the outside, not inside the hole, and in the case of surface polished specimens, the cracks occurred at the corners with the highest stress concentration. Table 7 shows the size of the initial fatigue crack generation site in the sample as-exposed.

3. Damage and EIFS correlation analysis of surface and fracture surface

EIFS (Equivalent Initial Flaw Size) values were derived based on the initial crack size observed at the fracture surface to determine the effect of surface damage in use on the remaining fatigue life of the structure. It is not an EIFS crack but refers to a specific critical size that exhibits crack growth behavior similar to cracks. In the present invention, Murakami's EIFS assumption is used. This method is applied to defects that may exist in casting materials, and is made by converting the cross-sectional area of the defects into equivalent initial cracks. Prior studies have predicted the remaining life of surface damage specimens present in P-3B aircraft components even within an error margin of approximately 10% less than other assumptions. Therefore, the life prediction using the same in the present invention was made, and the following equation was applied.

Using the above equation, the initial crack cross-sectional area at the fracture surface was converted to EIFS, and the correlation between the defect area measured at the fracture surface and the equivalent initial crack (EIFS) is shown in FIG. 3, where EIFS It was found that the size was approximately 30 to 60 µm. The above-described surface damage size measurement data were used to compare the size of the initial crack generation portion of the fatigue fracture surface with the corrosion damage present on the part surface. Corrosion damage present on the surface of the parts is on average 0.05 ~ 0.20mm2, the size of the initial crack generating portion of the fracture surface is 500 ~ 2500㎛ ² shows a difference of about 200 times.

표면 손상과 초기 피로 균열생성부의 200배 크기 차이 비율을 이용하여 위의 식의 A값에 표면손상크기를 이용한

를 대입하게 된다. 이 식을 이용하면 도 4와 같은 표면 손상 크기와 EIFS 간의 연관성으로 재해석될 수 있다.

Using the surface damage size and the surface damage size for the A value of the above equation,

Will be substituted. Using this equation can be reinterpreted as the correlation between the surface damage size and EIFS as shown in FIG.

4. Prediction of Residual Fatigue Life with EIFS

By applying the EIFS calculated as described above and the random load acting on the aircraft, it is possible to derive the residual fatigue life for the site. The reference load applied for the residual fatigue life analysis added 1.6 times the SN fatigue test load with constant load amplitude, and applied the FALSTAFF spectrum. 5 shows the results of crack growth analysis by deriving EIFS from 15% Wing Spar damaged specimens and surface polished specimens, respectively. As-polished represents the surface polished case, and Average defect age represents the result of crack growth analysis using EIFS calculated from damaged specimens. Looking at Figure 5, it can be seen that the fatigue life when damaged by long-term operation compared to the existing as-polished is reduced by about 50%. In fact, 15% Wing Spar-received parts were found to have no cracks or corrosion but were replaced after reaching the design life, and analysis and experiment show that the remaining life is about 50% .Wing 15% spar parts Applying the actual load spectrum imposed on, it is considered that more accurate life analysis results can be derived.

5. Prediction of Residual Fatigue Life Using Finite Element Analysis

식을

으로 변환할 수 있으며, 여기서 f(Kt)가 위의 유한요소해석으로부터 도출된 균열 끝단부에서의 응력집중계수이다.
In the present invention, the derivation of the EIFS size using the finite element analysis (FEM) as well as the derivation of the equation using the correlation between the surface corrosion damage and the fatigue fracture section tube. The finite element modeling of the infinite plate including the hole can derive the stress value as shown in FIG. 6. In general, when the stress applied to the plate is 115 MPa, the stress at the crack end is 4.5 times larger than this. Finite element modeling confirmed this. In this way, Murakami's

Expression

Where f (Kt) is the stress concentration factor at the crack tip derived from the finite element analysis above.

conclusion

In the present invention, the surface damage is observed quantitatively by observing the surface corrosion, defects or damage areas of the actual long-term operation of the aircraft structure, and also by performing a fatigue test using test specimens taken from the actual used parts, Correlation between corrosion damage and residual fatigue life was investigated. Through SN fatigue tests on aircraft parts that have been operated for a long time, quantitative evaluation of corrosion damage greatly reduces the fatigue life of aircraft structures, and the reduction of fatigue life of structures due to corrosion damage can be predicted properly using linear elastic fracture mechanics. It was confirmed that there is. By defining the relationship between damage observed at the surface and damage at the wavefront obtained through fatigue testing, a quantitative relationship between damage and residual fatigue life could be derived. In addition, by applying the finite element analysis method, the stress concentration coefficient could be effectively applied to the fatigue life analysis.

Although the embodiments of the present invention have been described above, various modifications may be made by those skilled in the art. Therefore, it should be understood that the present invention should not be construed as being limited to the above embodiments, but should be construed in accordance with the following claims.

Claims (6)

Selecting a target part of the long-term operating aircraft in order to predict the remaining fatigue life of the structure of the long-term operating aircraft (S100);
Database (S200) a surface corrosion form of the target component detached from the long-term operation aircraft;
Processing the SN fatigue test piece in the target part (S300),
Performing a fatigue test using the processed SN fatigue test piece (S400),
After the SN fatigue test, using the fatigue fracture surface to observe the initial fatigue crack generation site (S500),
And estimating fatigue life using crack propagation analysis code (S600). The method of claim 1,
In the step S100,
The target component is SN fatigue characteristics evaluation method according to the surface damage form of the long-term operation aircraft components, characterized in that the replacement parts to reach the design life without cracking during operation of the aircraft. The method of claim 1,
In the step (S300),
The SN fatigue test piece is a method of evaluating SN fatigue characteristics according to the surface damage pattern of the long-term operation aircraft component, characterized in that the plate fatigue specimen including a hole taken from the target component. The method of claim 1,
In the step S400,
The fatigue test,
Stress Ratio (Minimum Load / Maximum Load): 0.1
Cycle: 40㎐
Lab air
N _{run out} = 10 ^{7 A} method of evaluating SN fatigue characteristics according to the type of surface damage of a long-running aircraft component, characterized in that it is carried out under conditions of cycles (run-out refers to the service life of the specimen under any constant amplitude load). The method of claim 1,
In the step (S600),
Using crack propagation to predict the fatigue life,
The crack propagation equation applies NASGRO equation considering the lower limit stress intensity factor range and fracture toughness,
The NASGRO equation is

Where N is the number of working fatigue loads, a is the crack length, R is the stress ratio, K is the stress intensity factor, C, n, p and q are the material constants, and f is the crack open function method for evaluating SN fatigue characteristics according to surface damage patterns of long-term operation aircraft parts. The method of claim 1,
In the step S500,
EIFS (Equivalent Initial Flaw Size) refers to a specific critical size that is not a crack but exhibits crack growth behavior similar to a crack in order to determine the effect of surface damage occurring on the fatigue fracture surface on the residual fatigue life of the target part. SN fatigue characteristics evaluation method according to the form of damage to the surface of long-term operation aircraft components, characterized in that to derive.

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