A method for evaluating the strength recovery capability of aircraft local damage repair process

Abstract

The present invention provides a method for evaluating the strength recovery ability of a local damage repair process of an aircraft. A flat tensile test piece is manufactured, and a damage is prefabricated in the center of the test piece; the damage is equivalent to a crack; a correction factor Y1 related to the crack shape and a factor Y2 related to the plate configuration are obtained; the half-length a of the equivalent crack is measured, and the half-width b of the plate is measured; a tensile experiment is carried out; substituting into the remaining strength calculation formula of the structure before repair, the equivalent fracture toughness K is obtained D ; the test piece with the prefabricated damage is repaired by using the repair method to be evaluated; a tensile experiment is carried out again; substituting into the remaining strength calculation formula of the structure after repair, a configuration factor Y3 related to the repair process is obtained, and the reciprocal of which is defined as the strength recovery coefficient of a certain repair process. The present invention overcomes the deficiency that the existing evaluation method cannot fully consider the influence of the damage size ratio on the overall strength of the repaired structure, and can more accurately evaluate the strength recovery ability of a certain repair method.

Description

Technical Field

[0001] The invention relates to the technical field of aircraft damage repair, and in particular to a method for evaluating the strength recovery capability of aircraft local damage repair process. Background Art

[0002] Aircraft will inevitably encounter various damages during use. The most economical, effective and commonly used method is to perform in-situ repairs by properly selecting repair processes to restore the function of the damaged structure.

[0003] In the selection of aircraft repair technology, many factors need to be considered. Among them, the structural strength recovery ability is most closely related to the flight safety of the aircraft and is the most important indicator of concern. Therefore, how to evaluate the ability of a repair method to restore the aircraft structural strength is an important issue faced in the research and optimization of aircraft structural repair technology.

[0004] Since in many cases, aircraft structural damage is localized and small in scope, it is very important to evaluate the impact of local damage repair on the overall load-bearing capacity of the structure. However, this problem has not yet been satisfactorily resolved.

[0005] At present, the evaluation of aircraft repair process strength recovery rate is usually done by Figure 3 The tensile test specimen shown in the figure is used. That is, a through hole is first opened in the center of the plate to simulate damage. Then, the repair is carried out by gluing, riveting, or a combination of gluing and riveting. Figure 3 The tensile test is then carried out, and the strength of the repaired component is divided by the original strength of the plate before the hole is opened. The resulting percentage is defined as the structural strength recovery rate of the repair process.

[0006] This method is inaccurate because it does not consider the influence of damage size. For example, when the damage size is small, assuming that the damage size is so small that the structural strength loss rate is only 10%, at this time, no matter what repair process is used, the strength of the repaired structure will be almost more than 90% compared with the original structural strength. On the contrary, if the damage size is relatively large and the structural strength decreases more seriously, it is difficult to make the strength of the repaired structure reach 90% of the initial strength of the structure. In other words, the structural strength recovery rate obtained by the existing evaluation method is related to the size of the opening on the test piece, but the evaluation method fails to clearly point out the relationship between the recovery strength and the damage size. Therefore, the evaluation result is inaccurate and fails to reflect the true strength recovery ability of a repair method. Summary of the invention

[0007] In view of this, the present invention provides a method for evaluating the strength recovery capability of a local damage repair process of an aircraft.

[0008] In order to solve the above technical problems, the present invention provides a method for evaluating the strength recovery capability of aircraft local damage repair process, comprising the following steps:

[0009] A flat plate tensile test piece was made, and damage was preformed in the center of the test piece;

[0010] The damage is equivalent to a crack, and the correction factor Y related to the crack shape is obtained by consulting the stress intensity factor manual. 1 , factor Y related to the test panel configuration 2 ;

[0011] Measure half the length a of the equivalent crack and half the width b of the plate;

[0012] Carry out tensile test to obtain the ultimate strength σ of the damaged specimen C ;

[0013] Substitute the residual strength calculation formula of the structure before repair:

[0014]

[0015] Get the equivalent fracture strength K of the material D ;

[0016] Repair the damaged flat plate tensile specimen using the repair method to be evaluated;

[0017] The tensile test was carried out on the repaired specimen to obtain the ultimate strength σ′ of the repaired structure. C ;

[0018] Substitute the residual strength calculation formula of the repaired structure into:

[0019]

[0020] Get the configuration factor Y related to the repair process 3 , Y 3 It reflects the ability of the repair process to restore the strength of the damaged structure, and its reciprocal is defined as the strength recovery coefficient:

[0021]

[0022] Furthermore, the damage in the specimen is regarded as an equivalent crack, and the damage is trimmed into a regular, smooth circular hole, elliptical hole or rectangular hole during repair; when the damage is trimmed into a circular hole, a is the radius of the circular hole; when the damage is trimmed into an elliptical hole, a is half of the major axis of the ellipse; when the damage is trimmed into a rectangular hole, a is half of the length of the rectangular hole.

[0023] Furthermore, Y 2It is a factor related to the test panel configuration. For unrepaired bare panels, it is mainly related to the ratio of the length of the equivalent crack to the width of the panel and is calculated using the following formula:

[0024]

[0025] Furthermore, the equivalent fracture toughness of the material is determined by an experimental method using a circular hole defect specimen. Tensile tests are performed on specimens with different damage sizes to measure their ultimate strength σ C , the equivalent fracture strength is calculated using the following formula:

[0026]

[0027] Furthermore, the configuration factor Y related to the repair process 3 The tensile test of the repaired plate was carried out using a circular hole specimen with a hole radius of a to measure its ultimate strength σ′ C Then, use the following formula to calculate Y 3 :

[0028]

[0029] The beneficial effects of the above technical solution of the present invention are as follows:

[0030] The strength of an aircraft structure is related to the flight safety of the aircraft. If the strength of a repair method after repair cannot be accurately evaluated, it may pose a safety hazard to the use of the repaired aircraft, and the loss is difficult to estimate.

[0031] The evaluation method proposed in this invention overcomes the deficiency of existing evaluation methods that cannot fully consider the influence of damage size ratio on the overall strength of the repaired structure, and can more accurately evaluate the strength recovery ability of a repair method. The application of this achievement has an important guiding role in the research and development of emergency repair technology for aircraft structures, and is of great significance for ensuring the safe use of repaired aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0032] Figure 1 is the correction coefficient Y of the present invention 2 Relationship with the relative size a / b of the hole;

[0033] Figure 2 This is a residual strength analysis model diagram of the present invention;

[0034] Figure 3 Schematic diagram of the tensile test specimen for evaluating the aircraft riveting repair process of the present invention. DETAILED DESCRIPTION

[0035] To make the purpose, technical solution and advantages of the embodiments of the present invention clearer, the following will be combined with the appended drawings of the embodiments of the present invention. Figure 1-3 , the technical scheme of the embodiment of the present invention is clearly and completely described. Obviously, the described embodiment is a part of the embodiment of the present invention, not all of the embodiments. Based on the described embodiment of the present invention, all other embodiments obtained by ordinary technicians in this field belong to the scope of protection of the present invention.

[0036] This embodiment provides a method for evaluating the strength recovery capability of aircraft local damage repair process, based on fracture mechanics theory and structural residual strength analysis, and its basic principles are as follows:

[0037] The damage in the sample is considered as an equivalent crack. During repair, the damage is generally trimmed into a regular, smooth circular hole, elliptical hole or rectangular hole. In this embodiment, a circular hole is used, and treating it as an equivalent crack is a safer treatment method. When the damage is trimmed into a circular hole, a is the radius of the circular hole; when the damage is trimmed into an elliptical hole, a is half of the major axis of the ellipse; when the damage is trimmed into a rectangular hole, a is half of the length of the rectangular hole.

[0038] According to the fracture mechanics theory and the calculation method of residual strength of aircraft structures, the residual strength of damaged and unrepaired structures can be calculated using the following formula:

[0039]

[0040] In the formula, K D is the equivalent fracture strength of the material; a is half the length of the equivalent crack; Y 1 is a correction factor related to the crack shape, which can be determined by experiment (or directly queried from published literature); Y 2 It is a factor related to the test panel configuration. For unrepaired bare panels, it is mainly related to the ratio of the length of the equivalent crack to the width of the panel and can be calculated using the following formula:

[0041]

[0042] Where b is half the width of the plate, and the correction factor Y is 2 The relationship with the relative size a / b of the hole is as follows Figure 1 shown.

[0043] The equivalent fracture toughness of the material is determined by an experimental method using a circular hole defect specimen: a tensile test is performed on the specimens with different damage sizes (radius a) to measure their ultimate strength σ c , and then the equivalent breaking strength is calculated using the following formula.

[0044]

[0045] For the structure repaired by the process to be evaluated, the material of the component has not changed, the equivalent crack still exists, and the width of the component has not changed. It is only because of the addition of the repair patch that the configuration of the structure has changed. Therefore, the strength of the repaired structure can be calculated using the following formula:

[0046]

[0047] Where Y 3 is the configuration factor related to the repair process.

[0048] Configuration factor Y related to repair process 3 It can be determined by tensile test of the repaired component. In the test, a circular hole specimen with a hole radius of a is used to measure its ultimate strength σ c ′, then, use the following formula to calculate Y 3 .

[0049]

[0050] Y 3 It reflects the ability of the repair process to restore the strength of the damaged structure. The smaller the value, the stronger the strength recovery ability. In order to conform to the conventional thinking mode, a strength recovery coefficient is defined:

[0051]

[0052] The above method is only applicable to linear elastic characteristic segments.

[0053] For the formula

[0054] If we consider two extreme cases: one is that the damage size is zero, and the other is that the damage size is equal to the width of the plate. When the damage size is zero, the calculated structural residual strength is infinite; when the damage size is equal to the width of the plate, the calculated residual strength is still greater than zero. This is obviously not in line with the actual situation. In aircraft damage tolerance design, the above situation is treated as follows: Figure 2 The method shown is used to handle this.

[0055] like Figure 2 As shown, the strength corresponding to point A is the strength of the intact structure, and the damage size corresponding to point D is the size of the cross section of the structure. The AB and CD segments of the curve are transition segments, and BC is a linear elastic segment. The analysis method based on linear elastic fracture mechanics proposed in this patent is applicable to the BC segment. The above is a preferred embodiment of the present invention. It should be pointed out that for ordinary technicians in this technical field, several improvements and modifications can be made without departing from the principles of the present invention, and these improvements and modifications should also be regarded as the scope of protection of the present invention.

Claims

1. A method for evaluating the strength recovery capability of aircraft local damage repair process. Features: The steps include: A flat plate tensile test piece was made, and damage was preformed in the center of the test piece; The damage is equivalent to a crack, and the correction factor Y related to the crack shape is obtained by consulting the stress intensity factor manual. 1 , factor Y related to the test panel configuration 2 ; Measure half the length a of the equivalent crack and half the width b of the plate; Conduct tensile tests to obtain the residual strength σ of the structure before repair of the damaged test piece. C ; Substitute the residual strength calculation formula of the structure before repair: Get the equivalent fracture strength K of the material D ; Repair the damaged flat plate tensile specimen using the repair method to be evaluated; The tensile test was carried out on the repaired specimen to obtain the residual strength σ′ of the repaired structure. C ; Substitute the residual strength calculation formula of the repaired structure into: Get the configuration factor Y related to the repair process 3 , Y 3 It reflects the ability of the repair process to restore the strength of the damaged structure, and its reciprocal is defined as the strength recovery coefficient: Y 2 It is a factor related to the test panel configuration. For unrepaired bare panels, it is mainly related to the ratio of the length of the equivalent crack to the width of the panel and is calculated using the following formula: The damage in the specimen is regarded as an equivalent crack, and the damage is trimmed into a regular, smooth circular hole, elliptical hole or rectangular hole during repair. When the damage is trimmed into a circular hole, a is the radius of the circular hole; when the damage is trimmed into an elliptical hole, a is half of the major axis of the ellipse; when the damage is trimmed into a rectangular hole, a is half of the length of the rectangular hole.

2. The method for evaluating the strength recovery capability of the aircraft local damage repair process according to claim 1, Features: The equivalent fracture toughness of the material is determined by the test method using a circular hole defect specimen. A tensile test is performed on the specimens with different damage sizes to measure the residual strength σ of the structure before repair. C , the equivalent fracture strength is calculated using the following formula:

3. The method for evaluating the strength recovery capability of the aircraft local damage repair process according to claim 1, Features: Configuration factor Y related to repair process 3 The tensile test of the repaired plate was carried out using a circular hole specimen with a hole radius of a to measure the residual strength σ′ of the repaired structure. C Then, use the following formula to calculate Y 3 :

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