A laser shock strengthening method for multiple life extension repairs of aircraft structures

Abstract

This invention relates to the field of laser shock peening technology, and in particular to a laser shock peening method for multiple life extension repairs of aircraft structures. This method obtains the distribution function of fatigue crack initiation life and the residual stress relaxation model through structural simulation experiments. A safety factor is determined based on the initiation life distribution function, and a crack initiation life prediction model is established based on the residual stress relaxation model. By combining the predicted crack initiation life, the safety factor, and the fatigue load spectrum, the actual service interval for repairs of the aircraft structure is determined, and this process is applied to subsequent life extension repairs. This invention features simple principles and methods, comprehensive content, strong engineering guidance, and high operability. It can support the development of advanced aircraft structure fatigue resistance design and life extension repair schemes, and has significant engineering value and economic benefits.

Description

Technical Field

[0001] This invention relates to the field of laser shock peening technology, and in particular to a laser shock peening method for multiple life extension repairs of aircraft structures. Background Technology

[0002] Laser shock peening technology utilizes short pulses (on the order of nanoseconds) and high power density (>10). 9 W / cm 2 The laser beam passes through the transparent constraint layer and irradiates the surface of the metal material. The absorbing protective layer attached / coated to the surface absorbs the laser energy in a very short time, generating a high temperature (>10℃). 7 The plasma, at K and high pressure (>1 GPa), continues to absorb laser energy, generating a plasma shock wave (on the order of GPa) under the confinement of the confinement layer, which propagates into the interior of the metallic material. The force effect of the shock wave causes the material surface to undergo extremely high strain rates (10⁻⁶ K / L). 6 s -1 It produces plastic deformation and changes the surface microstructure, forming a residual compressive stress layer of a certain depth (mm level), thereby significantly improving properties such as fatigue resistance, stress corrosion resistance and wear resistance.

[0003] Research on laser shock peening technology and its strengthening mechanism has been extensive both domestically and internationally. However, for components that have been in service for a certain period, the residual compressive stress will relax under fatigue loads, and the original strengthening effect will gradually weaken with the extension of service time. According to relevant maintenance standards, they need to be returned to the factory for further strengthening and repair. However, there are very few research reports both domestically and internationally on how to use laser shock peening technology to carry out multiple life extension repairs, especially on how to determine the timing of the next laser shock peening.

[0004] Existing technologies have proposed solutions for determining the timing of a second laser shock peening repair, but these methods and techniques rely entirely on simulation models and numerical calculations. Furthermore, these solutions only address how to determine the timing of the second life extension repair, without discussing how to implement multiple life extension repairs, leaving a technological gap in guiding multiple life extension repairs of aircraft structures. Summary of the Invention

[0005] In view of the shortcomings of the prior art described above, the purpose of this invention is to provide a laser shock peening method for multiple life extension repairs of aircraft structures, which solves the problem of the time interval between multiple laser shock peening life extension repairs in the prior art.

[0006] To achieve the above and other related objectives, the present invention provides a laser shock peening method for repeated life extension repairs of aircraft structures, the method comprising:

[0007] A structural simulation component is fabricated, and the structural simulation component is subjected to laser shock hardening treatment;

[0008] The simulated structure is subjected to fatigue tests until it fractures, and the total fatigue life N0 of the simulated structure is obtained. Simultaneously, the distribution function N of the fatigue crack initiation life is derived by reverse engineering the fracture surface of the simulated structure. (x) ;

[0009] The distribution law of residual stress σ after laser shock strengthening treatment of the simulated structure and the relaxation model σ under fatigue load were obtained. (t) ;

[0010] Based on the combined results of fatigue tests and the residual stress relaxation model σ (t) A fatigue crack initiation life prediction model M based on residual stress parameters was established. (σ) And predict the fatigue crack initiation life N1 under the first laser shock strengthening treatment;

[0011] According to the distribution function N of the fatigue crack initiation life (x) The safety factor α (0 < α < 1) is determined by the function dispersion parameter;

[0012] The fatigue life N2' corresponding to the second laser shock strengthening treatment is obtained by multiplying the fatigue crack initiation life N1 and the safety factor α. The repair time interval in the actual service of the aircraft structure is obtained by back-calculating the fatigue load spectrum.

[0013] During the subsequent rework process, when performing the Nth laser shock peening treatment, the actual residual stress value after laser shock peening treatment and the fatigue crack initiation life prediction model M were used. (σ) The fatigue crack initiation life is predicted, and the next repair interval is determined based on the safety factor and fatigue load spectrum. The aircraft structure is then subjected to laser shock strengthening treatment according to the next repair interval.

[0014] Optionally, the fabrication of the structural simulation component and the laser shock hardening treatment of the structural simulation component include:

[0015] Based on the material types and dimensions of the aircraft structure, structural simulation parts were manufactured, and laser shock peening processes were designed and implemented.

[0016] Optionally, fatigue tests are performed on the simulated structure until it fractures to obtain the total fatigue life N0 of the simulated structure. Simultaneously, the distribution function N of the fatigue crack initiation life is obtained by reverse engineering the fracture surface of the simulated structure. (x) ,include:

[0017] Fatigue tests were conducted on the structural simulation components based on the fatigue load spectrum extracted from the characteristics of the aircraft structure.

[0018] Optionally, the distribution law of residual stress σ after laser shock strengthening treatment of the structural simulation part and the relaxation model σ under fatigue load are obtained. (t) ,include:

[0019] The distribution law of residual stress σ and the relaxation model σ under fatigue load were obtained by X-ray diffraction test and finite element simulation. (t) .

[0020] Optionally, the fatigue crack initiation life prediction model M based on residual stress parameters is... (σ) And predict the fatigue crack initiation life N1 under the first laser shock peening treatment, including:

[0021] The fatigue crack initiation life N1 is predicted based on the actual residual stress value after the first laser shock peening treatment.

[0022] As described above, the laser shock peening method for repeated life extension repair of aircraft structures according to the present invention has the following beneficial effects:

[0023] This scheme combines experimental testing and finite element simulation to determine the cycle of re-laser shock peening treatment, performing re-peening before crack initiation, thereby achieving extended service life repair through multiple laser shock peening treatments. This method is characterized by its simple principles, comprehensive content, strong engineering guidance, and high operability. It can support the development of fatigue resistance design and service life extension repair schemes for advanced aircraft structures, possessing significant engineering value and economic benefits. Attached Figure Description

[0024] Figure 1 This is a flowchart of an embodiment of the present invention. Detailed Implementation

[0025] The following specific examples illustrate the implementation of the present invention. Those skilled in the art can easily understand other advantages and effects of the present invention from the content disclosed in this specification. The present invention can also be implemented or applied through other different specific embodiments, and various details in this specification can also be modified or changed based on different viewpoints and applications without departing from the spirit of the present invention.

[0026] It should be noted that the illustrations provided in this embodiment are only schematic representations of the basic concept of the present invention. Therefore, the drawings only show components relevant to the present invention and are not drawn according to the actual number, shape, and size of components in implementation. In actual implementation, the form, quantity, and proportion of each component can be arbitrarily changed, and the component layout may be more complex. The structures, proportions, sizes, etc., depicted in the accompanying drawings are only for illustrative purposes to aid those skilled in the art and are not intended to limit the implementation conditions of the present invention. Therefore, they have no substantial technical significance. Any modifications to the structure, changes in proportions, or adjustments to size, without affecting the effects and objectives of the present invention, should still fall within the scope of the technical content disclosed in the present invention. Furthermore, the terms such as "upper," "lower," "left," "right," "middle," and "one" used in this specification are only for clarity of description and are not intended to limit the scope of the present invention. Changes or adjustments to their relative relationships, without substantially altering the technical content, should also be considered within the scope of the present invention.

[0027] Please see Figure 1 This embodiment provides a laser shock strengthening method for multiple life extension repairs of aircraft structures, the method comprising:

[0028] A structural simulation component is fabricated, and the structural simulation component is subjected to laser shock hardening treatment;

[0029] The simulated structure is subjected to fatigue tests until it fractures, and the total fatigue life N0 is obtained. The total fatigue life N0 includes the sum of the fatigue crack initiation life and the fatigue crack propagation life. Simultaneously, the distribution function N of the fatigue crack initiation life is obtained by reverse engineering the fracture surface of the simulated structure. (x) ;

[0030] The distribution law of residual stress σ after laser shock strengthening treatment of the simulated structure and the relaxation model σ under fatigue load were obtained. (t) ;

[0031] Based on the combined results of fatigue tests and the residual stress relaxation model σ (t) A fatigue crack initiation life prediction model M based on residual stress parameters was established. (σ) And predict the fatigue crack initiation life N1 under the first laser shock strengthening treatment;

[0032] According to the distribution function N of the fatigue crack initiation life (x) The safety factor α (0 < α < 1) is determined by the function dispersion parameter;

[0033] The fatigue life N2' corresponding to the second laser shock peening treatment is obtained by multiplying the fatigue crack initiation life N1 by the safety factor α. The actual repair interval of the aircraft structure in service is calculated by back-calculating the fatigue load spectrum. When performing the Nth laser shock peening treatment in subsequent repair processes, the actual residual stress value after laser shock peening treatment and the fatigue crack initiation life prediction model M are used as the basis for calculation. (σ) The fatigue crack initiation life is predicted, and the next repair interval is determined based on the safety factor and fatigue load spectrum. The aircraft structure is then subjected to laser shock strengthening treatment according to the next repair interval.

[0034] In one embodiment, the fabrication of the structural simulation component and the laser shock hardening treatment of the structural simulation component include:

[0035] Based on the material types and dimensions of the aircraft structure, structural simulation parts were manufactured, and laser shock peening processes were designed and implemented.

[0036] In one embodiment, fatigue tests are performed on the structural simulation component until it fractures to obtain the total fatigue life N0 of the structural simulation component. Simultaneously, the distribution function N of the fatigue crack initiation life is obtained by reverse engineering the fracture surface of the structural simulation component. (x) ,include:

[0037] Fatigue tests were conducted on the structural simulation components based on the fatigue load spectrum extracted from the characteristics of the aircraft structure.

[0038] In one embodiment, the distribution law of residual stress σ after laser shock strengthening treatment of the structural simulation component and the relaxation model σ under fatigue load are obtained. (t) ,include:

[0039] The distribution law of residual stress σ and the relaxation model σ under fatigue load were obtained by X-ray diffraction test and finite element simulation. (t) .

[0040] In one embodiment, the establishment of a fatigue crack initiation life prediction model M based on residual stress parameters (σ) And predict the fatigue crack initiation life N1 under the first laser shock peening treatment, including:

[0041] The fatigue crack initiation life N1 is predicted based on the actual residual stress value after the first laser shock peening treatment.

[0042] As described above, the laser shock peening method for repeated life extension repairs of aircraft structures in this embodiment has the following beneficial effects:

[0043] This embodiment combines experimental testing and finite element simulation to determine the cycle of subsequent laser shock peening treatment, performing re-peening before crack initiation, thereby achieving extended service life repair through multiple laser shock peening treatments. This method is characterized by its simple principles, comprehensive content, strong engineering guidance, and high operability. It can support the development of fatigue resistance design and service life extension repair schemes for advanced aircraft structures, possessing significant engineering value and economic benefits.

[0044] This embodiment obtains the distribution function of fatigue crack initiation life and the residual stress relaxation model through structural simulation tests. A safety factor is determined based on the initiation life distribution function, and a crack initiation life prediction model is established based on the residual stress relaxation model. By combining the predicted crack initiation life, the safety factor, and the fatigue load spectrum, the actual service life repair interval of the aircraft structure is determined, and this can be extrapolated to subsequent life extension repairs. This embodiment features simple principles and methods, comprehensive content, strong engineering guidance, and high operability. It can support the development of advanced aircraft structure fatigue resistance design and life extension repair schemes, and has significant engineering value and economic benefits.

[0045] The above embodiments are merely illustrative of the principles and effects of the present invention and are not intended to limit the invention. Any person skilled in the art can modify or alter the above embodiments without departing from the spirit and scope of the present invention. Therefore, all equivalent modifications or alterations made by those skilled in the art without departing from the spirit and technical concept disclosed in the present invention should still be covered by the claims of the present invention.

Claims

1. A laser shock peening method for repeated life extension repairs of aircraft structures, characterized in that, The method includes: A structural simulation component is fabricated, and the structural simulation component is subjected to laser shock peening treatment. The simulated structure is subjected to fatigue tests until it fractures, and the total fatigue life N0 of the simulated structure is obtained. Simultaneously, the distribution function N of the fatigue crack initiation life is derived by reverse engineering the fracture surface of the simulated structure. (x) ; The distribution law of residual stress σ after laser shock strengthening of the structural simulation part and the relaxation model σ under fatigue load were obtained. (t) ; Based on the combined results of fatigue tests and the residual stress relaxation model σ (t) A fatigue crack initiation life prediction model M based on residual stress parameters was established. (σ) And predict the fatigue crack initiation life N1 under the first laser shock strengthening treatment; According to the distribution function N of the fatigue crack initiation life (x) The safety factor α (0 < α < 1) is determined by the function dispersion parameter; The fatigue life N2' corresponding to the second laser shock strengthening treatment is obtained by multiplying the fatigue crack initiation life N1 and the safety factor α. The repair time interval in the actual service of the aircraft structure is obtained by back-calculating the fatigue load spectrum. During the subsequent rework process, when performing the Nth laser shock peening treatment, the actual residual stress value after laser shock peening treatment and the fatigue crack initiation life prediction model M were used. (σ) The fatigue crack initiation life is predicted, and the next repair interval is determined based on the safety factor and fatigue load spectrum. The aircraft structure is then subjected to laser shock strengthening treatment according to the next repair interval.

2. The laser shock peening method for multiple life extension repairs of aircraft structures according to claim 1, characterized in that: The process of fabricating a structural simulation component and subjecting the structural simulation component to laser shock hardening treatment includes: Based on the material types and dimensions of the aircraft structure, structural simulation parts were manufactured, and laser shock peening processes were designed and implemented.

3. The laser shock peening method for multiple life extension repairs of aircraft structures according to claim 1, characterized in that: The fatigue test is performed on the simulated structure until it fractures, obtaining the total fatigue life N0 of the simulated structure. Simultaneously, the distribution function N of the fatigue crack initiation life is obtained by reverse engineering the fracture surface of the simulated structure. (x) ,include: Fatigue tests were conducted on the structural simulation components based on the fatigue load spectrum extracted from the characteristics of the aircraft structure.

4. The laser shock peening method for repeated life extension repair of aircraft structures according to claim 1, characterized in that: The distribution law of residual stress σ after laser shock strengthening treatment of the structural simulation component and the relaxation model σ under fatigue load were obtained. (t) ,include: The distribution law of residual stress σ and the relaxation model σ under fatigue load were obtained by X-ray diffraction test and finite element simulation. (t) .

5. The laser shock peening method for repeated life extension repair of aircraft structures according to claim 1, characterized in that: The fatigue crack initiation life prediction model M based on residual stress parameters is established. (σ) And predict the fatigue crack initiation life N1 under the first laser shock peening treatment, including: The fatigue crack initiation life N1 is predicted based on the actual residual stress value after the first laser shock peening treatment.

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