CN112507457B - Method for evaluating calendar life of airplane structure - Google Patents

Method for evaluating calendar life of airplane structure Download PDF

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CN112507457B
CN112507457B CN202011432894.3A CN202011432894A CN112507457B CN 112507457 B CN112507457 B CN 112507457B CN 202011432894 A CN202011432894 A CN 202011432894A CN 112507457 B CN112507457 B CN 112507457B
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刘治国
李旭东
孙强
穆志韬
苏锋
张世禄
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Qingdao Campus of Naval Aviation University of PLA
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Abstract

The invention discloses an aircraft structure calendar life evaluation method, which comprises the following steps of S1: establishing a fatigue accumulated damage model of the aircraft material in a corrosive environment to obtain the corrosion equivalent damage of the aircraft material; s2: researching the change rule of the fatigue influence coefficient of the aircraft material along with the equivalent corrosion age in the corrosion environment to obtain the relation S3 between the corrosion fatigue influence coefficient of the aircraft material and the equivalent corrosion age: and evaluating the structural calendar life of the airplane according to the results obtained in the steps S1 and S2. The invention provides a fatigue damage accumulation model based on corrosion equivalent damage, the model error is small, and certain theoretical and practical values are achieved; the influence coefficient K of the corrosion fatigue is obtained T T change rule along with corrosion time is K T =1-0.5109 · exp (-0.241 · T); on the basis, a linear damage accumulation model based on corrosion equivalent damage and corrosion fatigue influence coefficients is provided for calculating the calendar life, and the estimation result has a certain reference value.

Description

Method for evaluating calendar life of airplane structure
Technical Field
The invention relates to the technical field of airplane service life estimation, in particular to an airplane structure calendar service life evaluation method.
Background
The number of landing times TTN, the number of flight hours TFH and the calendar life TCL of the aircraft constitute three major indicators of the life of the aircraft. Before the airplane is in service, rated values TTN, TFH and TCL of three indexes of the airplane are definitely given, and the actual criteria of the airplane for stopping flying and retirement are as follows: (TTN to-TTN) (TFH to-TFH) (TCL to-TCL) =0; the above equation shows that whichever of the three indicators reaches the rated value means the end of the useful life of the aircraft. From the technical route for determining the service life of the airplane, the service life of the number of landing times TTN or the number of flight hours TFH of the airplane is closely connected with the design idea of the airplane, and the calendar life is focused on the influence of the service environment.
The service environment of an aircraft includes a ground parking environment and an air flight environment. On one hand, airports are mostly distributed in coastal areas, the ground parking environment is high in temperature, high in humidity and long in salt mist occurrence time, and the airplane parking environment is quite severe under the action of industrial waste gas in coastal cities. In addition, the flying strength of the airplane in our army is low, the airplane is parked on the ground for about 97 percent of the time, so that most of the surface protective coating systems of the airplane structural parts are aged and peeled off, the base material is seriously corroded, and some parts, even main force-bearing parts, are seriously corroded. These seriously affect the calendar life of the aircraft. On the other hand, damage caused by corrosion fatigue in the air during flight is more serious than that caused by simple fatigue and corrosion damage, and the flight life is also reduced.
The above situation illustrates that under conditions of low annual flight intensity, the lead time, overhaul and overall life of the aircraft may still be controlled primarily by calendar life. At present, in the academic world, scholars at home and abroad make great efforts for scientifically determining the calendar life of an airplane structure, and the basic research approaches are roughly three: firstly, the research on corrosion fatigue is developed, and an S-N curve, an epsilon-Nf curve, a crack propagation rate curve and the like of different materials in different corrosion environments are given through a large number of corrosion fatigue tests; and secondly, the influence of the air weak corrosion environment is ignored, and the influence of the ground parking corrosion environment on the fatigue performance of the material is only researched. And thirdly, researching the influence of corrosion on the service life of the airplane structure, wherein the calendar service life of the metal part is considered to depend on the corrosion degree of the corrosion environment on the metal substrate, and at present, the research on the calendar service life in academic circles at home and abroad is one or the combination of the three basic researches.
However, no matter which damage mode is adopted to analyze the fatigue life, the damage calculation is finally carried out by using a fatigue accumulated damage theory, and whether the selection success of the accumulated damage theory directly relates to the precision of the fatigue life analysis, which can be called as 'missing milli centimetre and mu in thousand miles'. Over the past 100 years, researchers in various countries around the world have developed hundreds of fatigue cumulative damage theories for their specific problems, and the well-known Miner's linear cumulative damage theory has gained wide engineering application due to its simplicity. However, after considering the corrosive environment, how the damage is described, accumulated, what the damage threshold is, etc. need to be re-determined.
Disclosure of Invention
In view of the existing problems, the invention aims to provide an aircraft structure calendar life assessment method, which is used for assessing the structural calendar life of an aircraft by researching a fatigue accumulated damage rule in a corrosion environment and a change rule of a corrosion fatigue influence coefficient along with a calendar corrosion age.
In order to achieve the purpose, the technical scheme adopted by the invention is as follows:
a method for evaluating calendar life of an aircraft structure, comprising the steps of,
s1: establishing a fatigue accumulated damage model of the aircraft material in a corrosive environment to obtain the corrosion equivalent damage of the aircraft material;
s2: researching the change rule of the fatigue influence coefficient of the aircraft material along with the equivalent corrosion age in the corrosion environment to obtain the relation between the corrosion fatigue influence coefficient of the aircraft material and the equivalent corrosion age;
s3: and evaluating the structural calendar life of the airplane according to the results obtained in the step S1 and the step S2.
Further, the specific operation of step S1 includes the following steps,
s11: carrying out accelerated corrosion on the test piece to obtain corrosion pieces with different equivalent calendar years;
s12: static tensile test is carried out on the corroded parts with non-corrosion and different equivalent calendar years, and the tensile strength sigma of the material is determined b ；
S13: respectively carrying out comparison tests on an unwetted test piece and a corroded test piece which is subjected to pre-corrosion and corresponds to an equivalent calendar for one year under pure mechanical fatigue and corrosion fatigue conditions to determine the fatigue life of the test piece under different conditions;
s14: and establishing a fatigue accumulated damage model of the airplane in the corrosion environment according to the result of the step S13 to obtain the corrosion equivalent damage of the airplane material.
Further, in the step S11, the material of the test piece is a LY12CZ aluminum alloy, and the mass fraction of the chemical components of the material is as follows: cu:3.8-4.9%, mg:1.2-1.8%, mn:0.3-0.9%, fe:0.5%, si:0.5%, al: the balance;
the preparation method of the etching solution used for accelerating the etching in the step S1 comprises the following steps: preparing 5% NaCl solution with distilled water, adding small amount of 3-5% diluted H 2 SO 4 And the pH value of the solution reaches 4.0, the pH value is measured by adopting precise pH test paper in the test process, the pH value of the solution is kept within the range of 4.0-4.5, and the solution is replaced every 72 hours.
Further, the fatigue loading mode in the step S13 includes constant amplitude loading, high-low, low-high secondary loading and random spectrum loading;
in the two-stage loading test, a cyclic ratio n is firstly acted under the first-stage stress level 1 /N 1 Then to failure under a second level of stress; wherein the stress levels of high-low loading include: 0.65 sigma b -0.50σ b 、0.85σ b -0.50σ b And 0.85. Sigma b -0.65σ b (ii) a Stress levels for low-high loading include: 0.50 sigma b -0.65σ b 、0.50σ b -0.85σ b And 0.65 σ b -0.85σ b ；
The random spectrum loading test causes the loading stress to be 0.50 sigma respectively b 、0.65σ b 、0.85σ b The stress ratio is 0.06, the loading times are 1420, 608 and 20 cycles respectively, and 2048 cycle tests are randomly arranged by utilizing a multiplication remainder method to form a random block spectrum.
Further, the specific operation of step S14 includes the following steps,
s141: the acceleration effect of corrosion fatigue on pure mechanical fatigue is expressed using a corrosion fatigue influence coefficient, whereinCoefficient of influence of corrosion fatigue
Figure BDA0002827217640000031
In the formula, K (T) The corrosion fatigue influence coefficient is the calendar life T of the test piece; n is a radical of F(T) The corrosion fatigue life of the test piece after the equivalent corrosion age T is shown; n is a radical of (T) The test piece is subjected to equivalent corrosion age T and then subjected to pure mechanical fatigue life in a laboratory environment;
s142: the damage caused by one cycle of load working conditions in the general environment is
Figure BDA0002827217640000032
The corrosion damage caused by one cycle under the same working condition under the corrosion fatigue condition is
Figure BDA0002827217640000033
In the formula (I), the compound is shown in the specification,
Figure BDA0002827217640000034
coefficients related to load conditions and corrosive environments;
s143: the corrosion damage caused by corrosion is equivalent to corresponding fatigue damage according to the accumulated damage theory, namely the corrosion equivalent damage
Figure BDA0002827217640000035
In the formula, N 0 The life of the material in the general environment, N T The test piece is subjected to equivalent corrosion age T and then subjected to pure mechanical fatigue life in a laboratory environment;
s144: by adopting Miner fatigue accumulated damage theory, the corrosion equivalent damage is
Figure BDA0002827217640000036
Further, the specific operation steps of step S2 include,
s21: carrying out pure mechanical fatigue and corrosion fatigue tests on corrosion pieces subjected to corrosion of different equivalent ages in a general environment under random spectrum loading; fatigue loading stress of 0.50 sigma b 、0.65σ b And 0.85 σ b Stress ratio of 0.06, and loading times of 1420, 608 and 20 cycles respectively;
s22: carrying out pure mechanical fatigue and corrosion fatigue tests on corrosion parts subjected to corrosion in different equivalent years under a common environment by carrying out constant-amplitude spectrum loading; fatigue loading stress of 0.50 sigma b Stress ratio of 0.06;
s23: statistical analysis and verification of the pure mechanical fatigue and corrosion fatigue test data in the steps S21 and S22;
s24: analyzing corrosion equivalent damage data of the corrosion piece corroded by different equivalent ages to obtain a change curve equation of equivalent corrosion damage along with corrosion time T;
s25: and analyzing the corrosion fatigue influence coefficient data of the corrosion piece corroded by different equivalent years to obtain a change curve equation of the corrosion fatigue influence coefficient along with the corrosion time T.
Further, the specific operation of step S24 includes the steps of,
s241: fitting the corrosion equivalent damage D of the corrosion pieces corroded in different equivalent years with the equivalent corrosion year T to obtain D (T) = beta.T α ；
S242: for D (T) = beta.T α Transforming two equal-sign sides, and taking logarithms on the two sides to obtain a fitted regression line relational expression LnD (T) = Ln beta + alpha Ln (T);
s243: and (3) performing statistical analysis on the regression line obtained by fitting in the step S242 from the three aspects of variance statistics, curve slope and ordinate intercept to obtain a common equivalent corrosion damage fitting line Ln (D) = -0.9384+0.3403Ln (T) and a change curve equation D (T) = 0.3913T of equivalent corrosion damage along with the corrosion time T 0.3403 。
Further, the specific operation of step S25 includes the steps of,
s251: according to the formula
Figure BDA0002827217640000041
Calculating the corrosion fatigue influence coefficient K under different pre-corrosion years T of different load spectrums (T) ；
S252: fitting the corrosion fatigue influence coefficient and the corrosion age to obtain an exponential distribution equation K of the corrosion fatigue influence coefficient (T) ＝1-b 0 exp(b 1 T)；
S253: the exponential distribution equation in the step S252 is linearized to obtain ln (1-K) (T) ＝lnb 0 +b 1 T；
S254: the correlation influence of different loading spectrum types on the corrosion fatigue influence coefficient is verified through comparative analysis, the random spectrum and the constant-amplitude spectrum corrosion fatigue influence coefficient curve have no obvious difference, and the common corrosion fatigue influence coefficient curve is Ln (1-K) (T) ) Equation K for the curve of the fatigue coefficient of influence as a function of the corrosion time T, = -0.6716-0.241T T ＝1-0.5109*exp(-0.241*T)。
Further, the specific operation step of step S3 includes assuming that the aircraft has a life span under general circumstances
Figure BDA0002827217640000042
Hour, annual flying strength of delta N hours, flying life hours in corrosive environment of N ci (ii) a The method comprises the steps that a curve D (T) of corrosion equivalent damage D of a key part of an airplane in a parking environment along with equivalent corrosion age T is determined, and a curve K (T) of corrosion fatigue influence coefficients of the key part along with calendar age is determined in the flying process;
the number of hours of flight life of the aircraft in a corrosive environment N ci Is calculated by the formula
Figure BDA0002827217640000051
N ci ＝ΔN×T。
The beneficial effects of the invention are:
1. the invention provides a fatigue damage accumulation model based on corrosion equivalent damage, which has small model calculation error and certain theoretical and practical values;
2. in the invention, the relationship between the corrosion fatigue influence coefficient and the corrosion age limit of the airplane material is researched, and the trend that the corrosion fatigue influence coefficient shows an increase along with the increase of the corrosion age limit is further explainedThe essence is as follows: the corrosion fatigue influence coefficient is increased progressively along with the corrosion equivalent damage; and obtaining the influence coefficient K of corrosion fatigue T T change rule along with corrosion time is K T ＝1-0.5109·exp(-0.241·T)；
3. The invention provides a linear damage accumulation model based on corrosion equivalent damage and corrosion fatigue influence coefficients for calculating the calendar life, and the estimation of the calendar life of a key part of a certain airplane structure is carried out by using the linear damage accumulation model, and the estimation result has a certain reference value, so that the estimation method of the calendar life of the airplane structure can reflect the load-environment course actually suffered by the airplane structure more truly, has fewer parameters needing to be determined, and is convenient for engineering application.
Drawings
FIG. 1 is a flow chart of an aircraft structure calendar life assessment method of the present invention;
FIG. 2 is a flowchart of a test for establishing a fatigue cumulative damage model of an aircraft material in a corrosive environment according to the present invention;
FIG. 3 is a geometric dimension of a test piece used in the present invention;
FIG. 4 is an equivalent spectrum of accelerated corrosion test LY12CZ at the Cluster island airport of the present invention;
FIG. 5 shows the corrosion condition of the equivalent calendar year of the test piece of the present invention at 1 year;
FIG. 6 shows the 3-year corrosion status of the equivalent calendar year of the test piece in the present invention;
FIG. 7 shows the corrosion status of the equivalent calendar year of the test piece of the present invention at 5 years;
FIG. 8 shows the corrosion status of the equivalent calendar year of the test piece in the invention at 8 years;
FIG. 9 shows the corrosion status of the equivalent calendar year of the test piece of the present invention at 13 years;
FIG. 10 shows the stress level of 0.50 σ for the low-to-high loading of the present invention b -0.65σ b In time, the fatigue condition of the non-corroded test piece and the equivalent corrosion test piece for 1 year under pure mechanical fatigue and corrosion fatigue;
FIG. 11 shows the stress level of 0.50 σ for the low-to-high loading of the invention b -0.85σ b In time, the fatigue condition of the non-corroded test piece and the equivalent corrosion test piece for 1 year under pure mechanical fatigue and corrosion fatigue;
FIG. 12 shows the stress level of 0.65 σ for the low-to-high loading of the present invention b -0.85σ b In time, the fatigue condition of the non-corroded test piece and the equivalent corrosion test piece for 1 year under pure mechanical fatigue and corrosion fatigue;
FIG. 13 shows the stress level of the invention at high-low loading of 0.65 σ b -0.50σ b In the process, the fatigue conditions of the non-corroded test piece and the equivalent corrosion test piece for 1 year under pure mechanical fatigue and corrosion fatigue are met;
FIG. 14 shows the stress level of 0.85 σ for the high-low loading of the present invention b -0.50σ b In the process, the fatigue conditions of the non-corroded test piece and the equivalent corrosion test piece for 1 year under pure mechanical fatigue and corrosion fatigue are met;
FIG. 15 shows the stress level of 0.85 σ for the high-low loading of the present invention b -0.65σ b In the process, the fatigue conditions of the non-corroded test piece and the equivalent corrosion test piece for 1 year under pure mechanical fatigue and corrosion fatigue are met;
FIG. 16 is a schematic of a stochastic spectrum of fatigue loading in the present invention;
FIG. 17 is a graph showing the influence coefficient K of the graph processing software Origin on the corrosion fatigue under different pre-corrosion years T under the stochastic spectrum in the present invention (T) Performing linear fitting to obtain a linear fitting graph;
FIG. 18 shows that the graph processing software Origin is utilized to determine the influence coefficient K of the corrosion fatigue under different pre-corrosion age limits T under the constant-amplitude spectrum (T) Performing linear fitting to obtain a linear fitting graph;
FIG. 19 shows the C-T curve fitting results in the example of the present invention.
Detailed Description
In order to make those skilled in the art better understand the technical solution of the present invention, the following further describes the technical solution of the present invention with reference to the drawings and the embodiments.
As shown in fig. 1, a method for estimating calendar life of an aircraft structure includes the following steps,
step S1: establishing a fatigue accumulated damage model of the aircraft material in a corrosive environment to obtain the corrosion equivalent damage of the aircraft material;
the purpose of studying the corrosion fatigue cumulative damage law is to estimate the calendar life of the aircraft. And the calendar life of the airplane is short, more than ten years and more than 30 years, the complete simulation of the actual corrosion process is almost impossible, and the test is meaningless due to the overlong test period, so that a laboratory accelerated corrosion method is required to achieve the same corrosion effect as that of the airplane in service for several years in a shorter accelerated test period.
Specifically, the test procedure for establishing the fatigue accumulated damage model of the aircraft material in the corrosive environment is shown in fig. 2, and step S11: carrying out accelerated corrosion on the test piece to obtain corrosion pieces with different equivalent calendar years;
the test piece material is LY12CZ aluminum alloy, and the mass components of the chemical components of the material are as follows: cu:3.8-4.9%, mg:1.2-1.8%, mn:0.3-0.9%, fe:0.5%, si:0.5%, al: the allowance, the collective size of the test pieces, is shown in fig. 3.
The preparation method of the corrosion solution used for accelerating corrosion comprises the following steps: preparing 5% NaCl solution with distilled water, adding small amount of 3-5% diluted H 2 SO 4 And the pH value of the solution reaches 4.0, the pH value is measured by adopting precise pH test paper in the test process, the pH value of the solution is kept within the range of 4.0-4.5, and the solution is replaced every 72 hours.
According to a relational expression that fatigue performance of an airplane structure changes along with corrosion time and based on the viewpoint that the fatigue strength is equal when the corrosion damage is the same, an equivalent spectrum of accelerated corrosion environment of the cluster airport is determined according to a comparison test by using a corrosion equivalent principle taking the fatigue strength as a criterion, as shown in figure 4.
In the accelerated corrosion test, a ZJF-75G periodic infiltration corrosion test box is adopted, after accelerated corrosion of 1 year, 3 years, 5 years, 8 years and 13 years of equivalent calendar years is carried out on a test piece, the surface of a test section loses metallic luster, is in grey white, generates a large amount of black rusty spots, has no obvious corrosion products, and has basically the same phenomenon of each equivalent calendar year but different severity degrees, as shown in attached figures 5-9.
Further, step S12: static tensile test is carried out on corroded parts which are not corroded and have different equivalent calendar yearsEstablishing the tensile strength σ of the material b ；
Specifically, 3 specimens each of which had not been corroded and had equivalent calendar years of 1 year, 3 years, 5 years, 8 years and 13 years were subjected to static tensile tests. Its tensile strength sigma b The mean and coefficient of variation of (a) are shown in table 1 below.
TABLE 1 tensile Strength σ of Material for different calendar environments b Influence of (2)
Figure BDA0002827217640000071
As can be seen from Table 1, the static strength σ for equivalent calendar environments of 0 year, 1 year, 3 years, 5 years, 8 years and 13 years b The static strength values under the environment time of each equivalent calendar are relatively close, and the average value is approximate to sigma b ＝473Mpa。
Further, step S13: respectively carrying out comparison tests on an un-corroded test piece and a corroded test piece which is subjected to pre-corrosion and corresponds to an equivalent calendar for one year under pure mechanical fatigue and corrosion fatigue conditions, and determining the fatigue life of the test piece under different conditions;
all fatigue tests are carried out on a Material Test System 810 electro-hydraulic servo fatigue testing machine, control software is Basic Test Ware software provided by American MTS company, a loading waveform is a Sine wave, the waveform adopts PVC compensation, the loading frequency is 8-20 Hz, the dynamic load precision of the testing machine is +/-2%, and the static load precision is +/-1%; in order to simulate a corrosion environment during a corrosion fatigue test, a small box of the corrosion environment is added, the upper end of the small box is opened, the lower end of the small box is sealed, a saline solution with the pH value of 4-4.5 (a small amount of sulfuric acid is added) is immersed above a test section of a test piece, and the pH value of the solution is observed and kept to be about 4 in the test.
The fatigue loading modes comprise constant amplitude loading, high-low, low-high secondary loading and random spectrum loading, and all test working conditions are shown in the following table 2.
TABLE 2 fatigue test conditions
Figure BDA0002827217640000081
The fatigue life cycle numbers of the non-corroded test piece and the equivalent corrosion 1-year test piece under the conditions of pure mechanical fatigue and corrosion fatigue under the equal-amplitude loading are shown in the following table 3, the average value of the life obtained by the test is taken as the life value N of the test piece under the working condition, and the maximum working stress is 0.85 sigma respectively b ＝402Mpa、0.65σ b ＝307MPa、0.50σ b =236MPa, stress ratio 0.06.
Table 3 fatigue life test data under constant amplitude load (R = 0.06)
Figure BDA0002827217640000082
In the two-stage loading test, a cyclic ratio n is firstly acted under the first-stage stress level 1 /N 1 Then to failure under a second level of stress; wherein the stress levels of high-low loading include: 0.65 σ b -0.50σ b 、0.85σ b -0.50σ b And 0.85 σ b -0.65σ b (ii) a Stress levels for low-high loading include: 0.50 sigma b -0.65σ b 、0.50σ b -0.85σ b And 0.65 σ b -0.85σ b 。
Stress level of 0.50 sigma for low-to-high loading b -0.65σ b In the case of the non-corroded test piece and the equivalent corrosion test piece for 1 year, the fatigue conditions under pure mechanical fatigue and corrosion fatigue are shown in FIG. 10, and the stress level of the low-high loading is 0.50 sigma b -0.85σ b In time, the fatigue conditions of the non-corroded test piece and the equivalent corrosion 1 year test piece under pure mechanical fatigue and corrosion fatigue are shown in figure 11, and the stress level of the low-high loading is 0.65 sigma b -0.85σ b In the process, the fatigue conditions of the non-corroded test piece and the equivalent corrosion test piece for 1 year under pure mechanical fatigue and corrosion fatigue are shown in the attached drawing 12, and in the attached drawings 10-12, a dotted line is a Miner theoretical prediction curve; the chain line is a Manson model prediction curve; the dotted line is a ductile dissipation model prediction curve; the solid line is the data fit curve.
Under normal laboratory conditions, at low-high loading, due to the so-called "low-load exercise effect"
Figure BDA0002827217640000091
After the addition of "corrosive conditions", the test results for low-high loading show that:
1) Low-high loading of 0.50 σ whether un-corroded or corroded b -0.65σ b 、0.50σ b -0.85σ b And 0.65 σ b -0.85σ b In the mode, a fitting curve of pure mechanical fatigue data is similar to a fitting curve of corrosion fatigue data;
2) Except for 0.50 sigma b -0.85σ b Under the loading mode, the loading data points of the corroded test piece are not completely met, and the corroded test piece has the load data points under low-high loading
Figure BDA0002827217640000092
While the corroded test piece has
Figure BDA0002827217640000093
Fig. 13-15 show the results of high-low loading tests, showing,
1) High-low loading of 0.65 σ whether non-corroded test piece or pre-corroded test piece b -0.50σ b 、0.85σ b -0.50σ b And 0.85 σ b -0.65σ b In the mode, a fitting curve of pure mechanical fatigue data is similar to a fitting curve of corrosion fatigue data;
2) Under high-low loading to non-corroded test piece
Figure BDA0002827217640000094
For pre-corroded test pieces, except for the loading mode, the method comprises the following steps
Figure BDA0002827217640000095
In conclusion, if the pre-corrosion effect is not achieved, only the corrosion fatigue effect is achieved, and the test result of the two-stage loading fatigue accumulated damage under the conventional condition cannot be changed from the quality aspect; and the pre-corrosion condition changes the test result of two-stage loading fatigue accumulated damage under the conventional condition from the quality.
Further, the random spectrum loading test is carried out so that the loading stress is respectively 0.50 sigma b 、0.65σ b 、0.85σ b The stress ratio is 0.06, the loading times are 1420, 608 and 20 cycles respectively, and 2048 cycle tests are randomly arranged by using a multiplication remainder method to form a random block spectrum. The data values of pure mechanical fatigue and corrosion fatigue life under random spectral loading for the non-corroded test piece and the test piece with an equivalent calendar life of 1 year are listed in the following table 4.
TABLE 4 fatigue test data under random spectra
Figure BDA0002827217640000101
Table 4 was converted to injury data calculated by Miner's theory, and the results are shown in Table 5.
TABLE 5 fatigue test cumulative damage data under random spectra
Figure BDA0002827217640000102
As can be seen from table 5, the damage D calculated using the Miner's theory was near 1 and was less than 1 for both the un-corroded and corroded specimens, and for both the pure mechanical and corrosion fatigue, when the random spectra were loaded.
Further, step S14: and establishing a fatigue accumulated damage model of the airplane in the corrosive environment according to the result of the step S13.
Specifically, S141: the accelerated effect of the corrosion fatigue on the pure mechanical fatigue is expressed by using the corrosion fatigue influence coefficient, and as the change rule of the corrosion fatigue influence coefficient of the LY12CZ aluminum alloy test piece in a single acidic corrosion medium with 3.5 percent of NaCl and 4-4.5 of PH value is mainly researched, the corrosion fatigue influence coefficient is defined
Figure BDA0002827217640000103
In the formula, K (T) The corrosion fatigue influence coefficient is the calendar life T of the test piece; n is a radical of hydrogen F(T) The corrosion fatigue life of the test piece after the equivalent corrosion age T is shown; n is a radical of (T) The test piece is subjected to equivalent corrosion age T and then subjected to pure mechanical fatigue life in a laboratory environment;
s142: the damage caused by one cycle of load working conditions in the general environment is
Figure BDA0002827217640000104
The corrosion damage caused by one cycle under the same working condition under the corrosion fatigue condition is
Figure BDA0002827217640000105
In the formula (I), the compound is shown in the specification,
Figure BDA0002827217640000106
is a coefficient related to the load condition and the corrosion environment;
under the condition of secondary loading, the corrosion fatigue cumulative damage model based on the corrosion fatigue influence coefficient is verified, and it can be obtained that whether the test piece is an un-corroded test piece or a corroded test piece, whether the test piece is high-low loaded or low-high loaded, the change rule of the pure mechanical fatigue data has high similarity with the change rule of the corrosion fatigue data, and therefore the correctness of the corrosion fatigue cumulative damage model based on the corrosion fatigue influence coefficient is qualitatively explained. The corrosion fatigue accumulated damage model based on the corrosion fatigue influence coefficient is quantitatively verified by adopting a corrected Miner theory, and the result shows that the error of converting the corrosion fatigue into the pure mechanical fatigue by using the corrosion fatigue influence coefficient is smaller as well, and the prediction effect is better.
S143: the corrosion damage caused by corrosion is equivalent to corresponding fatigue damage according to the accumulated damage theory, namely the corrosion equivalent damage
Figure BDA0002827217640000111
In the formula, N 0 The life of the material in the general environment, N T Is a test piecePure mechanical fatigue life in a laboratory environment after an equivalent corrosion age T;
s144: by adopting Miner fatigue accumulated damage theory, the corrosion equivalent damage is
Figure BDA0002827217640000112
Under the condition of secondary loading, verifying a damage accumulation model based on corrosion equivalent damage, and obtaining: the prediction error of the fatigue accumulated damage model based on the corrosion equivalent damage, which is provided by the invention, on the pure mechanical fatigue secondary loading of the corrosion test piece and the error predicted by using the Miner theory are good or bad along with the difference of the first-stage loading circulation ratio, but the prediction error of the fatigue accumulated damage model based on the corrosion equivalent damage is smaller than the error predicted by using the Miner theory on the whole. Therefore, the present model can be considered to be effective. From a comparison of the prediction error of the fatigue cumulative damage model based on corrosion equivalent damage with "+" of the error predicted using the Miner theory, the fatigue cumulative damage model based on corrosion equivalent damage substantially agrees with the Miner theoretical prediction error. And the prediction effect of the Miner theory on the random spectrum loading is better due to the error offset effect when the Miner theory is used for loading the random spectrum. Therefore, the fatigue accumulated damage model based on the corrosion equivalent damage can be deduced to better predict the random spectrum loading result.
Further, step S2: researching the change rule of the fatigue influence coefficient of the aircraft material along with the equivalent corrosion age in the corrosion environment to obtain the relationship between the corrosion fatigue influence coefficient of the aircraft material and the equivalent corrosion age;
specifically, step S21: carrying out pure mechanical fatigue and corrosion fatigue tests on corrosion pieces subjected to corrosion of different equivalent ages in a general environment under random spectrum loading; the random spectrum is schematically shown in FIG. 16. Fatigue loading stress of 0.50 sigma b 、0.65σ b And 0.85. Sigma b Stress ratio of 0.06, and loading times of 1420, 608 and 20 cycles respectively; and randomly arranging the 2048 cycles by utilizing a multiplication remainder method to form a random block spectrum.
Step S22: for different passesCarrying out pure mechanical fatigue and corrosion fatigue tests on the year-dependent corroded corrosion piece in a common environment with constant-amplitude spectrum loading; fatigue loading stress of 0.50 sigma b Stress ratio of 0.06;
the results of the pure mechanical fatigue and corrosion fatigue tests of the test pieces under random spectral loading and constant amplitude spectral loading are shown in tables 6 and 7.
TABLE 6 pure mechanical fatigue and Corrosion fatigue test results under random Spectrum loading
Figure BDA0002827217640000121
TABLE 7 results of pure mechanical fatigue and corrosion fatigue tests under loading of equi-amplitude spectra
Figure BDA0002827217640000122
S23: statistical analysis and verification of the pure mechanical fatigue and corrosion fatigue test data in the step S21 and the step S22;
specifically, the distribution function of the fatigue life data is tested by using a W test method, and the result shows that the distribution of the fatigue life still better follows the lognormal distribution.
Taking the common logarithm of the service life 1gN as a sample to carry out inspection, and respectively carrying out inspection on the variance of the fatigue life under various conditions by using an F inspection; the difference condition of the mean values is detected by using a t-test method, and the result shows that the mean values of four different working conditions, namely equal-amplitude spectrum fatigue, equal-amplitude spectrum corrosion fatigue, random spectrum fatigue and random spectrum corrosion fatigue, under different corrosion ages are obviously detected. It is shown that after different equivalent years of pre-corrosion, the test piece has different degrees of corrosion damage, which causes significant change of service life.
To what extent the fatigue life data obtained by fatigue testing can represent the overall useful life of the material, it needs to be examined by interval estimation. When the mean life of the test sample is used as the estimation quantity of the mean life of the parent body, the confidence coefficient of gamma is used, and the relative error does not exceed +/-delta:
Figure BDA0002827217640000131
in the formula, t γ The t distribution quantity under the confidence coefficient gamma; s is the standard deviation of the test sample;
Figure BDA0002827217640000134
log mean life for test samples; n is the number of test samples. The relative error is basically within 5% at a 95% confidence level, which shows that the test result can well represent the overall life of the material.
S24: analyzing corrosion equivalent damage data of the corrosion piece corroded by different equivalent years;
specifically, S241: fitting the corrosion equivalent damage D of the corrosion pieces corroded in different equivalent years with the equivalent corrosion year T to obtain D (T) = beta.T α ；
According to the formula
Figure BDA0002827217640000133
The equivalent corrosion damage data of the random spectrum and the constant-amplitude spectrum pure mechanical fatigue are calculated and shown in table 8.
TABLE 8 Corrosion equivalent Damage data
Equivalent corrosion damage 1 year 3 years old 5 years old 8 years old 13 years old
Random spectral loading 0.2964 0.6000 0.6575 0.7148 0.8225
Constant amplitude spectrum loading 0.4283 0.6905 0.7990 0.8554 0.9015
The fitting comparison shows that the linear relation between the corrosion equivalent damage D and the natural logarithm ln (T) of the equivalent corrosion age T is the best, but the initial condition of D (0) =0 cannot be met. Therefore, better linear fitting of the natural logarithm Ln (D) and the natural logarithm Ln (T) is adopted. Namely: d (T) = β · T α 。
S242: for D (T) = beta.T α Transforming two equal-sign sides, and taking logarithms on the two sides to obtain a fitted regression line relational expression LnD (T) = Ln beta + alpha Ln (T);
the results after the fitting are shown in table 9 below.
TABLE 9 regression line parameters
Figure BDA0002827217640000141
At the significance level α, the critical value γ of the linear correlation coefficient satisfying the hypothesis distribution can be calculated by the following formula
Figure BDA0002827217640000142
In the formula, t (n-2) represents a t distribution having a degree of freedom of n-2. Substituting the formula can obtain the gamma of alpha =0.05 c =0.8054. Comparison fitting knotAs a result, the linear fitting effect of the natural logarithm Ln (T) is better.
S243: the regression line obtained by fitting in the step S242 is subjected to statistical analysis from the three aspects of variance statistics, curve slope and ordinate intercept,
and (3) variance statistical comparison:
Figure BDA0002827217640000143
looking up F distribution table to obtain F 0.975 (3,3) =15.44 since F < F 0.975 (3, 3), so that the two are considered to have no significant difference, have homogeneous variance and total standard deviation of
Figure BDA0002827217640000144
Statistical analysis of the slope of the curve:
Figure BDA0002827217640000145
looking up t distribution table to obtain t 0.975 (6) =5.82, since | t | < t 0.975 (6) Therefore, the two are considered to have no significant difference, and the two are combined to have a common slope of
Figure BDA0002827217640000146
Vertical coordinate intercept statistical analysis:
Figure BDA0002827217640000147
looking up t distribution table to obtain t 0.975 (6) =5.82, since | t | < t 0.975 (6) Therefore, the two are considered to have no significant difference, and the two are combined into a common ordinate intercept of
Figure BDA0002827217640000151
Therefore, the corrosion equivalent damage curves of the random spectrum and the constant-amplitude spectrum have no significant difference, and a common equivalent corrosion damage fitting straight line Ln (D) = -0.9384+0.3403Ln (T) exists, so that the change curve equation D (T) = 0.3913T of the equivalent corrosion damage along with the corrosion time T 0.3403 。
It follows that corrosion equivalent damage is closely related to the corrosion environment and corrosion time, and not to the load spectrum. Under certain corrosive environments, the corrosion equivalent damage increases with the age of corrosion.
Further, step S25: analyzing the corrosion fatigue influence coefficient data of the corrosion piece corroded by different equivalent years;
specifically, S251: according to the formula
Figure BDA0002827217640000152
Calculating the corrosion fatigue influence coefficient K under different pre-corrosion years T of different load spectrums (T) The results are shown in Table 10 below.
TABLE 10 influence coefficient of corrosion fatigue K under different pre-corrosion age T of different load spectra (T)
Figure BDA0002827217640000153
S252: the corrosion fatigue influence coefficient increases with the increase of the corrosion age, and when the corrosion reaches a certain degree, the influence of the corrosion fatigue is very small and hardly influenced, so that the corrosion fatigue influence coefficient curve approaches to 1 when the T reaches a certain value. Fitting the corrosion fatigue influence coefficient and the corrosion age to obtain an exponential distribution equation K of the corrosion fatigue influence coefficient (T) ＝1-b 0 exp(b 1 T)；
S253: the exponential distribution equation in step S252 is linearized to obtain ln (1-K) (T) ＝ln b 0 +b 1 T；
The data in table 10 were fitted with a straight line using the graphics processing software Origin to obtain a straight line fit for both spectral patterns, as shown in fig. 17 and 18, and the fit equation and correlation coefficient are shown in table 11.
TABLE 11 regression line parameters
Figure BDA0002827217640000161
Using the formula
Figure BDA0002827217640000162
Calculating a critical value gamma of the linear correlation coefficient under the condition of alpha =0.05 c =0.9000. And comparing the fitting results to obtain: the correlation of the straight line fitting is quite good, and the selected type equation can well fit the test result.
S254: the correlation influence of different loading spectrum types on the corrosion fatigue influence coefficient is verified through comparative analysis, the obtained random spectrum and constant-amplitude spectrum corrosion fatigue influence coefficient curves have no obvious difference, and the common corrosion fatigue influence coefficient curve is Ln (1-K) (T) ) = -0.6716-0.241T, i.e. K T ＝1-0.5109*exp(-0.241*T)。
Specifically, statistical analysis is carried out from three aspects of variance statistics, curve slope and ordinate intercept:
and (3) variance statistical comparison:
Figure BDA0002827217640000163
looking up F distribution table to obtain F 0.975 (2,2) =39, since F < F 0.975 (1, 1), so the two are considered to have no significant difference, have the same variance and the total standard deviation of
Figure BDA0002827217640000164
Statistical analysis of the slope of the curve:
Figure BDA0002827217640000165
looking up the distribution table to obtain t 0.975 (4) =2.7764, since | t | < t 0.975 (4) Therefore, the two are considered to have no significant difference, and the two are combined to have a common slope of
Figure BDA0002827217640000166
Vertical coordinate intercept statistical analysis:
Figure BDA0002827217640000167
looking up t distribution table to obtain t 0.975 (4) =2.7764, since | t | < t 0.975 (4) Therefore, the two are considered to have no significant difference, and the two are combinedCombined to a common ordinate intercept of
Figure BDA0002827217640000168
Therefore, the obtained random spectrum and the constant-amplitude spectrum have no obvious difference on the corrosion fatigue influence coefficient curve, and the common corrosion fatigue influence coefficient curve is Ln (1-K) (T) ) = 0.6716-0.241 × t, and if there is no significant difference in the curve between the random spectrum and the constant-amplitude spectrum, the curve can be considered to be substantially independent of the form of the load spectrum, and K is therefore T ＝1-0.5109*exp(-0.241*T)。
Further, step S3: and evaluating the structural calendar life of the airplane according to the results obtained in the step S1 and the step S2.
The corrosion fatigue influence coefficient is constant for test pieces in the same initial state (mainly meaning having the same corrosion damage), and the corrosion fatigue damage equivalent can be converted into pure mechanical fatigue damage by using the coefficient. And the previous research shows that when the random spectrum is loaded, the damage accumulation error is smaller according to the linear Miner theory, and the random spectrum loading method is within the acceptable range of engineering application. Therefore, the corrosion damage in the whole parking process of the airplane can be equivalent to damage xi 1 Then the corrosion fatigue damage in each flight process is converted into pure mechanical fatigue damage through a corrosion fatigue influence coefficient, and the damage is linearly accumulated to xi regardless of the loading sequence effect 2 When damage xi 1 +ξ 2 When the threshold value 1 is reached, the aircraft is considered to be up to life.
Assuming that the aircraft has a life span of
Figure BDA0002827217640000171
Hour, annual flying strength of delta N hours, flying life hours in corrosive environment of N ci (ii) a The method comprises the steps that a curve D (T) of corrosion equivalent damage D of a key part of an airplane in a parking environment along with equivalent corrosion age T is determined, and a curve K (T) of corrosion fatigue influence coefficients of the key part along with calendar age is determined in the flying process;
the number of hours of flight life of the aircraft in a corrosive environment is N ci Is calculated by the formula
Figure BDA0002827217640000172
N ci ＝ΔN×T。
The embodiment is as follows:
taking a wing main beam of a typical model airplane in China as an example, the key dangerous part is 2 holes of a lower flange after the airplane is navigated. The conclusion of the service life determination under the general environment of the model is as follows: when the aircraft is turned firstly for 1200 flight hours, the structure is only checked and is not repaired; the second repair interval is 1000 flight hours, and the overhaul life is 2200 flight hours; the post-repair life is 800 flight hours and the total life is 3000 flight hours.
The simulated test piece with the protective coating at the position is adopted, the test piece material is a 30CrMnSiNi2A forging, the form is a tensile test piece with two phi 6 bolt holes, the thickness is 6mm, the width of the working section is 20mm, LY12 aluminum sheets with the thickness of 5mm are attached to the working section, and the tensile test piece is connected through bolts. The protection system of test piece is: after phosphorization, H06-2 iron red primer is coated, and then H04-2 steel ash enamel is coated. The physical conditions of the test piece such as material state, connection form, aperture, edge distance, thickness and other geometric dimensions, processing technology, loading form, protective layer and the like well simulate the real situation of the 2-hole position of the lower flange of the wing girder after navigation.
And carrying out accelerated corrosion test on the simulated test piece, wherein an accelerated test environment spectrum consists of two parts and is implemented by a periodic infiltration corrosion test box.
(1) Soaking in an acidic NaCl solution: 5% NaCl solution to add a little diluted H2SO4, pH =4-4.5, solution temperature 40 + -2 ℃.
(2) The test piece is dried by irradiating the test piece with a far infrared lamp in humid air at the temperature of 40 ℃ and RH = (95 +/-5)% and the position and power of the far infrared lamp are adjusted so that the test piece can be dried before being immersed in the solution.
One accelerated spectrum period is 30min, soaking for 7.5min, and standing for 22.5min.
And (3) taking the maximum load in the load spectrum before the test piece, and performing pre-stretching twice at a certain temperature.
And after the accelerated corrosion test is finished, performing a spectrum fatigue test in a general environment, wherein the load spectrum adopts a fatigue load spectrum of the airplane and a nominal stress spectrum formed by 1g of stress of each load state at a dangerous part of a flange at the root of the main beam. The spectrum has 83.7 flight hours as a fundamental cycle, and includes 15 mission profiles, 114 flight landing and landing, 7 load states, and 1228 load cycles. Due to the presence of the asymmetric loading condition, a period is actually constituted with 167.4 flight hours, and it is necessary to apply 6gDYD once after each period, and 7gDYD once after 12 fundamental periods (the fundamental period in which 7gDYD is applied does not apply 6gDYD load any more).
The experimental load spectrum is described by the following formula: p = k · n y Sigma (1 g) F, wherein k is a load adjustment coefficient, so that the median life of the non-corrosive test piece in a laboratory environment is 10000-20000 flight hours, and the medium life is obtained by probing the stress level; n is y Is an overload spectrum; sigma (1 g) is stress of 1g in each load state at dangerous positions of a lower flange of a main girder of the airplane wing; f is the net area of the cross section of the simulated test piece hole of the wing girder of the airplane, and the nominal value is 14 multiplied by 6=84mm 2 。
The pre-corrosion fatigue tests were carried out in 5 groups (corresponding to 0 year, 5 years, 10 years, 20 years, 30 years) under the above-mentioned environmental spectrum and load spectrum, and the test results and treatment parameters are shown in table 12.
TABLE 12 Pre-corrosion fatigue Life
T j (year) N 0j (hours of flight) Standard deviation of logarithmic life C j
0 10278 0.10 1
5 9163 0.13 0.89152
10 8595 0.13 0.83625
20 7573 0.098 0.73682
30 6841 0.061 0.6656
The above experimental data sets have a homogeneity of variance, taking the common significance α = 0.05. The C-T curve form is converted into linearity, fitting is carried out according to the least square method, and the fitting result of the C-T curve is shown in figure 19. The relation obtained by fitting is
C＝1.0-0.038704*T 0.635070 Correlation coefficient r =0.999
C＝exp(-0.035810*T 0.712801 ) Correlation coefficient r =0.999.
Further, the K value was determined in an air environment spectrum consisting of room temperature atmosphere, humid air, 3.5% of NaCl neutral salt spray and acid salt spray (3.5% 2 ) The four single media were composed, and the specific compositions are shown in table 13. Four groups of wing girder simulation test pieces are adopted to respectively carry out a spectrum-loaded fatigue test of a certain type of airplane under the four single medium environmentsThe percentage of each medium and the median corrosion fatigue life for a single medium are shown in Table 13.
TABLE 13 airborne Environment Spectrum composition
Media i y i N i0 k i
Ambient atmosphere 87.5 11933 1
Moist air 10 9135 0.762
Neutral salt spray 1.25 5286 0.424
Acid salt spray 1.25 5648 0.471
From Table 13It can be derived that,
Figure BDA0002827217640000191
typical annual flight intensity is 80, 100 flight hours, fatigue life under general environment is 800, 1000, 1200, 2200, 3000 flight hours, and the data are used to evaluate the fatigue life according to the method in "calendar life system evaluation technology of airplane structure" (nephrite-containing, lieyuhai, etc.; beijing: aeronautical industry press, 2004-138) (hereinafter referred to as "prior art"), and the results are shown in table 14.
TABLE 14 prior art fatigue Life assessment results
Figure BDA0002827217640000192
Flight life hours N of an aircraft in a corrosive environment using the invention ci Is calculated by
Figure BDA0002827217640000193
N ci The fatigue life of the aircraft was calculated by = Δ N × T, and the results are shown in table 15 below.
TABLE 15 evaluation results of the method of the present invention on the fatigue life of an airplane
Figure BDA0002827217640000201
From the above calculation, it can be seen that the life value obtained by using the life evaluation method based on the damage accumulation model of corrosion equivalent damage in the present invention is shorter than the life value given in the prior art, which is more safe, but the relative error is less than 10%, which can be used as a reference for life evaluation.
The foregoing shows and describes the general principles, principal features, and advantages of the invention. It will be understood by those skilled in the art that the present invention is not limited to the embodiments described above, which are described in the specification and illustrated only to illustrate the principle of the present invention, but that various changes and modifications may be made therein without departing from the spirit and scope of the present invention, which fall within the scope of the invention as claimed. The scope of the invention is defined by the appended claims and equivalents thereof.

Claims (1)

1. A method for evaluating calendar life of an aircraft structure, comprising the steps of,
s1: establishing a fatigue accumulated damage model of the aircraft material in a corrosive environment to obtain the corrosion equivalent damage of the aircraft material;
s2: researching the change rule of the fatigue influence coefficient of the aircraft material along with the equivalent corrosion age in the corrosion environment to obtain the relationship between the corrosion fatigue influence coefficient of the aircraft material and the equivalent corrosion age;
s3: evaluating the structural calendar life of the aircraft according to the results obtained in the step S1 and the step S2;
wherein, the specific operation of the step S1 comprises the following steps,
s11: carrying out accelerated corrosion on the test piece to obtain corrosion pieces with different equivalent calendar years;
s12: static tensile test is carried out on the corroded parts with non-corrosion and different equivalent calendar years, and the tensile strength sigma of the material is determined b ；
S13: respectively carrying out comparison tests on an un-corroded test piece and a corroded test piece which is subjected to pre-corrosion and corresponds to an equivalent calendar for one year under pure mechanical fatigue and corrosion fatigue conditions, and determining the fatigue life of the test piece under different conditions;
s14: establishing a fatigue accumulated damage model of the airplane in the corrosion environment according to the result of the step S13 to obtain the corrosion equivalent damage of the airplane material;
in the step S11, the test piece material is LY12CZ aluminum alloy, and the mass fraction of the chemical components of the material is as follows: cu:3.8-4.9%, mg:1.2-1.8%, mn:0.3-0.9%, fe:0.5%, si:0.5%, al: the balance;
the preparation method of the corrosion solution used for the accelerated corrosion in the step S1 comprises the following steps: preparing 5% NaCl solution with distilled water, adding small amount of 3-5% diluted H 2 SO 4 The pH value of the solution reaches 4.0, the pH value is measured by adopting precise pH test paper in the test process, the pH value of the solution is kept within the range of 4.0-4.5, and the solution is replaced every 72 hours;
the fatigue loading mode in the step S13 comprises constant amplitude loading, high-low, low-high secondary loading and random spectrum loading;
in the two-stage loading test, a cyclic ratio n is firstly acted under the first-stage stress level 1 /N 1 Then to failure under the second level stress level; wherein the stress levels of high-low loading include: 0.65 sigma b -0.50σ b 、0.85σ b -0.50σ b And 0.85. Sigma b -0.65σ b (ii) a Stress levels for low-to-high loading include: 0.50 sigma b -0.65σ b 、0.50σ b -0.85σ b And 0.65 σ b -0.85σ b ；
The random spectrum loading test causes the loading stress to be 0.50 sigma respectively b 、0.65σ b 、0.85σ b The stress ratio is 0.06, the loading times are 1420, 608 and 20 cycles respectively, and 2048 cycle tests are randomly arranged by utilizing a multiplication remainder method to form a random block spectrum;
the specific operation of step S14 includes the following steps,
s141: representing the accelerating effect of corrosion fatigue on pure mechanical fatigue using a corrosion fatigue influencing factor, wherein the corrosion fatigue influencing factor
Figure FDA0003872797780000021
In the formula, K (T) The corrosion fatigue influence coefficient is the calendar life T of the test piece; n is a radical of F(T) The corrosion fatigue life of the test piece after the equivalent corrosion age T; n is a radical of (T) The test piece is subjected to equivalent corrosion age T and then subjected to pure mechanical fatigue life in a laboratory environment;
s142: the damage caused by one cycle of load working conditions in the general environment is
Figure FDA0003872797780000022
Then the corrosion fatigue strip under the same working conditionOne cycle under the part causes corrosion damage to
Figure FDA0003872797780000023
In the formula (I), the compound is shown in the specification,
Figure FDA0003872797780000024
coefficients related to load conditions and corrosive environments;
s143: the corrosion damage caused by corrosion is equivalent to corresponding fatigue damage according to the accumulated damage theory, namely the corrosion equivalent damage
Figure FDA0003872797780000025
In the formula, N 0 The life of the material in the general environment, N T The test piece has a pure mechanical fatigue life in a laboratory environment after the equivalent corrosion age T;
s144: by adopting Miner fatigue accumulated damage theory, the corrosion equivalent damage is
Figure FDA0003872797780000031
The specific operation steps of step S2 include,
s21: carrying out pure mechanical fatigue and corrosion fatigue tests on corrosion pieces subjected to corrosion of different equivalent ages in a general environment under random spectrum loading; fatigue loading stress of 0.50 sigma b 、0.65σ b And 0.85. Sigma b Stress ratio of 0.06, and loading times of 1420, 608 and 20 cycles respectively;
s22: carrying out pure mechanical fatigue and corrosion fatigue tests on corrosion parts subjected to corrosion in different equivalent years under a common environment by carrying out constant-amplitude spectrum loading; fatigue loading stress of 0.50 sigma b Stress ratio of 0.06;
s23: statistical analysis and verification of the pure mechanical fatigue and corrosion fatigue test data in the step S21 and the step S22;
s24: analyzing the corrosion equivalent damage data of the corrosion piece corroded by different equivalent ages to obtain a change curve equation of equivalent corrosion damage along with corrosion time T;
s25: analyzing the data of the corrosion fatigue influence coefficients of the corrosion pieces corroded by different equivalent ages to obtain a change curve equation of the corrosion fatigue influence coefficients along with the corrosion time T;
the specific operation of step S24 includes the following steps,
s241: fitting the corrosion equivalent damage D of the corrosion pieces corroded in different equivalent years with the equivalent corrosion age T to obtain D (T) = beta.T α ；
S242: for D (T) = beta.T α Transforming two sides with equal sign, and taking logarithm on the two sides to obtain a fitted regression line relational expression LnD (T) = Ln beta + alpha Ln (T);
s243: and (3) performing statistical analysis on the regression line obtained by fitting in the step S242 from the three aspects of variance statistics, curve slope and ordinate intercept to obtain a common equivalent corrosion damage fitting line Ln (D) = -0.9384+0.3403Ln (T) and a curve equation D (T) = 0.3913T of the change of the equivalent corrosion damage along with the corrosion time T 0.3403 ；
The specific operation of step S25 includes the steps of,
s251: according to the formula
Figure FDA0003872797780000041
Calculating the corrosion fatigue influence coefficient K under different pre-corrosion years T of different load spectrums (T) ；
S252: fitting the corrosion fatigue influence coefficient and the corrosion age to obtain an exponential distribution equation K (T) =1-b of the corrosion fatigue influence coefficient 0 exp(b 1 T)；
S253: the exponential distribution equation in step S252 is linearized to obtain ln (1-K (T)) = ln b 0 +b 1 T；
S254: the correlation influence of different loading spectrum types on the corrosion fatigue influence coefficient is verified through comparative analysis, the obtained random spectrum and constant-amplitude spectrum corrosion fatigue influence coefficient curves have no obvious difference, and the common corrosion fatigue influence coefficient curve is Ln (1-K) (T) ) Equation K of the curve of the fatigue influence coefficient as a function of the corrosion time T = -0.6716-0.241T T ＝1-0.5109*exp(-0.241*T)；
The specific operation of step S3 includes assuming that the aircraft has a life span in a typical environment of
Figure FDA0003872797780000042
Hour, annual flying strength of DeltaN hour, and flying life of N hours in corrosive environment ci (ii) a The method comprises the steps that a curve D (T) of corrosion equivalent damage D of a key part of an aircraft along with the change of equivalent corrosion age T in a parking environment is determined, and a curve K (T) of corrosion fatigue influence coefficient of the key part along with the change of calendar age in the flying process is determined;
the number of hours of flight life of the aircraft in a corrosive environment N ci Is calculated by the formula
Figure FDA0003872797780000043
N ci ＝ΔN×T。
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CN104318128A (en) * 2014-11-18 2015-01-28 中国人民解放军空军工程大学 Method for determining calendar safe life of airplane structure protection system
CN104316457A (en) * 2014-11-18 2015-01-28 中国人民解放军空军工程大学 Method for determining reliability of calendar life of airplane structure protection system

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