Load spectrum simplification method for aircraft construction crack extension test

A technology for crack propagation test and aircraft structure, applied in the field of load spectrum simplification for crack propagation test of aircraft structure, can solve problems such as increasing development cost, restricting aircraft development time, long fatigue test cycle, etc., achieving large engineering application value and shortening the test. The effect of cycle and method is simple and easy to implement

Active Publication Date: 2014-09-10

CHINA AIRPLANT STRENGTH RES INST

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Problems solved by technology

However, in the prior art aviation fatigue/crack growth test, due to the complexity of the crack growth of aviation metal materials, the fati...

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Abstract

The invention discloses a load spectrum simplification method for an aircraft construction crack extension test. The load spectrum simplification method comprises the following steps: (1) determining the structural type of a crack extension test piece and a material containing a crack part, and determining the fatigue crack extension threshold value deltaKth of the material; (2) determining a stress intensity factor of a crack tip of the material; (3) estimating the crack length in the crack extension test and processing in a segmenting manner; (4) deleting the stress cycle lower than the fatigue crack extension threshold value in a load spectrum according to the segmentation condition to form a load spectrum corresponding to the lengths of the segments of the crack, and finally applying the load spectrum in the crack extension test. The load spectrum simplification method disclosed by the invention is used for determining the cark tip stress intensity factor value of a crack-containing structure and simplifying the load spectrum for the crack extension test in the segmenting manner in combination with the fatigue crack extension threshold value deltaKth of the material, so that the load spectrum can be directly applied in the aircraft construction crack extension test, and the method is simple and feasible and is capable of effectively shortening the test period.

Application Domain

Special data processing applicationsStrength properties

Technology Topic

AirplaneLoad spectrum +7

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[0031] The present invention will be described in further detail below by means of specific embodiments:

[0032] see figure 1 , which is the simplification method of the load spectrum for the crack growth test of the aircraft structure of the present invention-the segmental processing relationship diagram of the typical crack growth curve. The main characteristic parameters of the load spectrum simplification method for the aircraft structure crack growth test of the present invention include the crack growth threshold ΔK th , crack tip stress intensity factor K, half-crack length a, load spectrum.

[0033] During the whole process of crack growth, based on the idea of low load interception combined with DK th , K and a cut off the load spectrum in sections, by figure 1 It can be seen that the load spectrum loaded at each stage is different.

[0034] see figure 2 , which is a flow chart of the load spectrum simplification method for the aircraft structure crack growth test of the present invention, and the concrete process steps of the present invention are:

[0035] Step 1: Determine the structural form of the crack growth test piece and the material of the cracked part, and determine the fatigue crack growth threshold ΔK of the material th;

[0036] In the specific embodiment, for different materials, the fatigue crack growth threshold value ΔK can be measured according to the small test piece test th , for commonly used materials, it can be obtained by checking the "Stress Intensity Factor Handbook" published by Science Press in 1993, while for emerging materials, it needs to be measured by small test piece test.

[0037] Step 2: Determine the stress intensity factor at the tip of the structural crack;

[0038] The solution of the stress intensity factor at the crack tip of the cracked structure is a crucial step in this paper, which directly affects the accuracy and error of subsequent tests. The general function expression of the stress intensity factor is formula (1), while the expression of the stress intensity factor of a typical structure is formula (2):

[0039] K=f(β,σ,a) (1)

[0040] K = β πa - - - ( 2 )

[0041] In the formula:

[0042] σ—loading stress;

[0043] a—is the half-crack length;

[0044] β—parameters related to structural configuration parameters and crack length, referred to as configuration parameters;

[0045] It should be pointed out that the determination of the configuration parameter β is the most important thing, and the usual method is to check the "Stress Intensity Factor Handbook" according to the specific structure; for those that are not in the handbook or have a more complicated structure, usually The finite element method is used to solve the problem, and the finite element method often uses an indirect method to obtain the stress intensity factor, such as calculating the J integral and the opening displacement of the crack tip. Specific issues require specific analysis. For example, for a finite-width plate with cracks at the edge of the center hole, the expression of its configuration factor β is formula (3); and for a typical fuselage panel structure, if its frame is cracked, the expression of its configuration factor β The formula is formula (4); and for complex structures, such as wing-body joints, if cracks appear on the surface, the stress intensity factor needs to be calculated by finite element software, such as the J integral in Abaqus.

[0046] β = sec πa w - - - ( 3 )

[0047] β=β 1 beta 2 beta 3 … (4)

[0048] In the formula:

[0049] w—board width;

[0050] beta 1 , β 2 , β 3 — It is a parameter related to the structure, for details, please refer to the "Damage Tolerance Design Manual for Structural Durability";

[0051] The crack tip stress intensity factor K value calculated in step 2 is for comparison with the fatigue crack growth threshold, and since the K value is related to the crack length, K here is dynamic.

[0052] Step 3: Estimate the critical crack length in the crack growth test:

[0053] For general metal structures, in the crack growth test, the determination of the critical crack length includes the net section yield method, the fracture toughness criterion, the R curve criterion and the COD criterion, etc., and one or several of them can be selected to obtain the correct The critical crack length of a specific structure a c.

[0054]Step 4: Segment the crack length. The division of the crack length is closely related to the accuracy of the follow-up test. Therefore, the crack growth curve (a-N curve) of similar materials and structures should be used for reference when crack growth is segmented; the principle adopted here is:

[0055] a) Crack initiation stage and steady state propagation stage, multi-segmented;

[0056] b) Rapid expansion phase, divided into one segment;

[0057] The specific number of segments depends on the actual structure. The more segments, the less time will be saved in the process of crack growth. However, the load spectrum calls will take more time, which should be weighed. Here, a finite-width plate with cracks at the edge of the central hole is taken as an example for a brief description. Combining formula (2) and formula (3), and bringing in the threshold value of fatigue crack growth, formula (5) can be obtained. It can be seen that the plate width w and fatigue crack growth threshold ΔK in formula (5) th are constant, the half-crack length a is inversely proportional to the truncated stress value σ, and an important conclusion can be drawn: the more segments, the more small loads are truncated.

[0058] The segmentation method can be determined according to the standard of the cut-off load, such as determining the size of the cut-off load first, that is, σ i , and then calculate the corresponding crack length a by formula (5) i , then corresponding to a i When , all load spectrums less than σ i The stress values can be deleted, which simplifies the test load spectrum and shortens the test time.

[0059] Δ K th = σ πa · sec πa w - - - ( 5 )

[0060] Step 5: Cut off the load spectrum to form a series of new load spectra;

[0061] According to the segmentation situation in step 4, combined with the stress intensity factor calculation formula in step 2, the corresponding maximum stress intensity factor value under each crack length can be calculated, and the stress value in the load spectrum is calculated using the fatigue crack growth threshold Cut off, the specific method is as described in step 4, determine the corresponding half-crack length a i , the cut-off stress σ is calculated by formula (5), then the value smaller than the stress σ in the load spectrum can be deleted, and the magnitude and position of other stresses in the load spectrum remain unchanged. Finally, a series of test load spectra are formed. During the test, the crack length of the structure is observed in time. When the crack length reaches the length a used in the previous calculation i , select the corresponding load spectrum here for the test.

[0062] The load spectrum simplification method for the aircraft structure crack growth test of the present invention proposes a test loading method through the load spectrum segmentation processing technology, which can achieve the effect of shortening the test cycle, and the processing process is simple, strong in operability, and strong in effectiveness. Simple and easy to implement, it is very suitable for the crack growth test of aircraft structure (metal), and has great engineering application value.

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Classification and recommendation of technical efficacy words

Shorten the test cycle
Simple process

Load spectrum simplification method for aircraft construction crack extension test

AI-Extracted Technical Summary

Problems solved by technology

Abstract

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Example Embodiment

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Description & Claims & Application Information

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