A method of repairing damage to a composite structure on an aircraft

Abstract

The application belongs to the technical field of composite material structure damage repair on an airplane, and particularly relates to a composite material structure damage repair method on an airplane, which integrates an internal, quantitative quality and safety judgment system, can scientifically evaluate internal defects formed in a repair process, and thus ensures that repair of the structure can achieve an expected safety target.

Description

Technical Field

[0001] This application belongs to the field of aircraft composite material structure damage repair technology, specifically relating to a method for repairing damage to aircraft composite material structures. Background Technology

[0002] Composite materials are increasingly used in the main load-bearing structures of aircraft due to their high specific strength, high specific stiffness, and excellent designability. However, during service, composite structures are susceptible to various types of damage, such as delamination, perforation, and cracking, under loads such as impact and fatigue. To ensure flight safety and structural integrity, timely and reliable repair of such damage is essential.

[0003] Existing methods for repairing composite material structures mostly focus on the repair scheme itself. A prominent issue is how to scientifically verify whether the structure meets safety requirements after field repair. In view of this, this application is proposed.

[0004] This application is made in view of the aforementioned technical deficiencies. Summary of the Invention

[0005] The purpose of this application is to provide a method for repairing damage to composite material structures on aircraft, and a method for assessing damage to composite material structures on aircraft. It integrates a built-in, quantitative quality and safety assessment system, which can scientifically assess internal defects formed during the repair process, thereby ensuring that the repair of the structure can achieve the expected safety goals.

[0006] The technical solution of this application is:

[0007] A method for repairing damage to composite material structures on aircraft, comprising:

[0008] Step 1: Determine the damaged area of ​​the structure, query the minimum safety margin RF0 of the damaged area, and the key factors affecting the safety margin, and then remove the damage.

[0009] Step 2: Design a repair plan for the structure;

[0010] Step 3: Prepare repair test pieces for the damaged area according to the repair plan, and conduct ultrasonic non-destructive testing, mechanical property testing of key influencing factors, and metallographic testing on the repair test pieces for the damaged area.

[0011] Step 4: Fabricate calibration test pieces for the damaged area according to the materials and processes of the original structure, and conduct metallographic testing, ultrasonic non-destructive testing, and mechanical property tests on the calibration test pieces for the damaged area, as well as tests on key influencing factors.

[0012] Step 5: Calculate the echo intensity loss Δdb1 of the repair test piece relative to the calibration test piece in the damaged area, and the key influencing factor, the mechanical property reduction coefficient K.

[0013] Step 6: Using the minimum safety margin RF0 of the damaged area and the mechanical performance reduction coefficient K of the key influencing factor, calculate the repair safety margin RF of the damaged area and determine whether the repair plan meets the requirements.

[0014] Step 7: Repair the damaged areas of the structure according to the repair plan;

[0015] Step 8: Perform ultrasonic non-destructive testing on the structure, calculate the echo intensity loss Δdb2 of the damaged area relative to the undamaged area, compare it with Δdb1, and judge the repair effect on the damaged area.

[0016] According to at least one embodiment of this application, in the above-described method for repairing damage to composite material structures on aircraft, in step one, ultrasonic non-destructive testing is performed on the structure, and the damaged area is determined by the echo intensity.

[0017] According to at least one embodiment of this application, in the above-described method for repairing damage to composite material structures on aircraft, step two involves designing a repair scheme for the structure, including repair materials and repair processes.

[0018] According to at least one embodiment of this application, in the above-mentioned method for repairing damage to composite material structures on aircraft, in step three, multiple repair test pieces for damaged areas are made according to the repair plan, and the echo intensity db1 of the ultrasonic non-destructive testing of the repair test pieces for damaged areas is taken as the average of the maximum and minimum values.

[0019] The mechanical property test results of the key influencing factors of the repair test piece in the damaged area are recorded as F1.

[0020] According to at least one embodiment of this application, in the above-described method for repairing damage to composite material structures on aircraft, in step three, metallographic testing is performed at the locations of the maximum and minimum echo intensities of the repair test piece in the damaged area.

[0021] According to at least one embodiment of this application, in the above-described method for repairing damage to composite material structures on aircraft, in step three, if the metallographic test shows an average porosity greater than 3%, the repair plan is considered to be unacceptable and the repair plan is improved.

[0022] According to at least one embodiment of this application, in the above-mentioned method for repairing damage to composite material structures on aircraft, in step four, multiple damage area comparison test pieces are made according to the materials and processes of the original structure, the process allowance of each damage area comparison test piece is cut off, metallographic testing is performed, and the damage area comparison test piece with the smallest porosity obtained from the metallographic testing is selected as the damage area calibration test piece.

[0023] According to at least one embodiment of this application, in the above-described method for repairing damage to composite material structures on aircraft, in step four, the porosity obtained by metallographic testing of the damaged area calibration test piece is less than 2%.

[0024] According to at least one embodiment of this application, in the above-described method for repairing damage to composite material structures on aircraft, in step four, the echo intensity db2 of the ultrasonic nondestructive testing of the damaged area calibration test piece is taken as the average value of multiple locations.

[0025] The mechanical property test results of the key influencing factors of the damaged area calibration test piece are recorded as F2.

[0026] According to at least one embodiment of this application, in the above-described method for repairing damage to composite material structures on aircraft, step five, calculating the echo intensity loss Δdb1 of the repair test specimen in the damaged area relative to the calibration test specimen in the damaged area, specifically involves:

[0027] △db1 = db1 - db2.

[0028] According to at least one embodiment of this application, in the above-described method for repairing damage to composite material structures on aircraft, step five involves calculating the mechanical property reduction coefficient K of the key influencing factor, specifically as follows:

[0029] K = F1 / F2.

[0030] According to at least one embodiment of this application, in the above-described method for repairing damage to composite material structures on aircraft, step six, calculating the repair safety margin RF of the damaged area, specifically involves:

[0031] RF = K × (RF0 + 1) - 1.

[0032] According to at least one embodiment of this application, in the above-described method for repairing damage to composite material structures on aircraft, in step six, if the repair safety margin RF of the damaged area is less than 0, the repair plan is considered to be unacceptable, and the repair plan is improved.

[0033] According to at least one embodiment of this application, in the above-described method for repairing damage to composite material structures on aircraft, in step eight, the echo intensity db3 of the ultrasonic nondestructive testing of the structural damage area is taken as the average value of multiple locations.

[0034] According to at least one embodiment of this application, in the above-described method for repairing damage to composite material structures on aircraft, in step eight, the echo intensity db4 of the ultrasonic nondestructive testing of the undamaged area of ​​the structure is taken as the average value of multiple locations.

[0035] According to at least one embodiment of this application, in the above-described method for repairing damage to composite material structures on aircraft, step eight, calculating the echo intensity loss Δdb2 of the damaged area relative to the undamaged area, specifically involves:

[0036] △db2 = db3 - db4.

[0037] According to at least one embodiment of this application, in the above-described method for repairing damage to composite material structures on aircraft, in step eight, if Δdb2 > αΔdb1, it is considered that the repair effect on the damaged area is not good, and the damaged area of ​​the structure is repaired again according to the repair plan, where α is the redundancy coefficient.

[0038] According to at least one embodiment of this application, in the above-described method for repairing damage to composite material structures on aircraft, in step eight, the redundancy factor is taken as 1 to 1.05. Attached Figure Description

[0039] Figure 1 This is a schematic diagram of a method for repairing damage to composite material structures on aircraft provided in an embodiment of this application;

[0040] Figure 2 This is a schematic diagram of the composite material repair patch provided in the embodiments of this application.

[0041] To better illustrate this embodiment, some content in the accompanying drawings may be omitted, enlarged, or reduced. They are for illustrative purposes only and should not be construed as limiting the scope of this application. Detailed Implementation

[0042] To make the technical solution and advantages of this application clearer, the technical solution of this application will be described in a clearer and more complete manner below with reference to the accompanying drawings. It should be understood that the specific embodiments described herein are only some embodiments of this application, and are only used to explain this application, not to limit this application. It should be noted that, for ease of description, only the parts related to this application are shown in the accompanying drawings, and other related parts can be referred to the general design.

[0043] Furthermore, unless otherwise defined, the technical or scientific terms used in this application description shall have the ordinary meaning understood by one of ordinary skill in the art to which this application pertains. The word "comprising" as used in this application description indicates that the concept preceding the word encompasses the concepts listed following the word and their equivalents, without excluding other related concepts.

[0044] Furthermore, the terms indicating location used in the description of this application are only used to indicate relative directions or positional relationships. When the absolute position of the described object changes, its relative positional relationship may also change accordingly. It should also be noted that, unless otherwise explicitly specified and limited, terms such as "installation" and "connection" used in the description of this application should be interpreted broadly. For example, a connection can be a fixed connection or a detachable connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium. Those skilled in the art can understand its specific meaning in this application according to the specific circumstances.

[0045] A method for repairing damage to composite material structures on aircraft, such as Figure 1 As shown.

[0046] Step 1: Determine the damaged area of ​​the structure, query the minimum safety margin RF0 of the damaged area, and the key factors affecting the safety margin, and then remove the damage.

[0047] By performing ultrasonic non-destructive testing on the structure, the damaged area can be identified and determined through the echo intensity, and the type of damage can be further determined.

[0048] By consulting technical materials to determine the minimum safety margin RF0 specified for the damaged area, and analyzing and identifying the key influencing factors affecting the safety margin of the damaged area, targeted mechanical performance tests can be designed.

[0049] The minimum safety margin RF0 is calculated as follows: RF0 = Allowable value / Load value - 1.

[0050] Step 2: Design a repair plan for the structure.

[0051] The repair plan includes repair materials and repair processes that take into account the field repair and construction environment.

[0052] Step 3: Prepare a repair test piece for the damaged area according to the repair plan, and conduct ultrasonic non-destructive testing, mechanical property testing of key influencing factors, and metallographic testing on the repair test piece for the damaged area.

[0053] Multiple repair test pieces for damaged areas were made according to the repair plan. The echo intensity db1 of the ultrasonic non-destructive testing of the repair test pieces for damaged areas was taken as the average of the maximum and minimum values.

[0054] The mechanical property test results of the key influencing factors of the repair test piece in the damaged area are recorded as F1.

[0055] Metallographic examination is performed at the locations of the maximum and minimum echo intensity of the repair test piece in the damaged area. If the average porosity obtained by metallographic examination is greater than 3%, the repair plan is considered to be unacceptable. The repair plan is then improved and step three is repeated.

[0056] Step 4: Fabricate a calibration test piece for the damaged area according to the materials and processes of the original structure, and conduct metallographic testing, ultrasonic non-destructive testing, and mechanical property tests on the calibration test piece for the damaged area, as well as tests on key influencing factors.

[0057] Multiple comparative test specimens of damaged areas were fabricated according to the original structure's materials and processes. The process allowance of each comparative test specimen of damaged areas was cut off, and metallographic testing was performed. The comparative test specimen of the damaged area with the lowest porosity obtained from the metallographic test was selected as the damaged area calibration test specimen. At the same time, it was required that the porosity of the damaged area calibration test specimen obtained from the metallographic test be less than 2%.

[0058] The echo intensity db2 of the ultrasonic nondestructive testing of the damaged area calibration test piece is taken as the average value of multiple locations.

[0059] The mechanical property test results of the key influencing factors of the damaged area calibration test piece are recorded as F2.

[0060] Step 5: Calculate the echo intensity loss Δdb1 of the repair test piece relative to the calibration test piece in the damaged area, and the mechanical property reduction coefficient K, a key influencing factor.

[0061] △db1 = db1 - db2.

[0062] K = F1 / F2.

[0063] Step 6: Using the minimum safety margin RF0 of the damaged area and the mechanical performance reduction coefficient K of the key influencing factor, calculate the repair safety margin RF of the damaged area, and determine whether the repair plan meets the requirements.

[0064] RF = K × Allowable value / Load value - 1 = K × (RF0 + 1) - 1.

[0065] If the repair safety margin RF of the damaged area is less than 0, the repair plan is considered to be unacceptable. The repair plan should be improved and steps three, five, and six should be repeated.

[0066] Step 7: Repair the damaged areas of the structure according to the repair plan.

[0067] Step 8: Perform ultrasonic non-destructive testing on the structure, calculate the echo intensity loss Δdb2 of the damaged area relative to the undamaged area, compare it with Δdb1, and judge the repair effect on the damaged area.

[0068] The echo intensity db3 of ultrasonic nondestructive testing in the structural damage area is taken as the average value of multiple locations.

[0069] The echo intensity db4 of the undamaged area of ​​the structure is taken as the average value of multiple locations.

[0070] △db2 = db3 - db4.

[0071] If △db2>α△db1, then the repair effect on the damaged area is considered to be unsatisfactory. The damaged area of ​​the structure is repaired again according to the repair plan, and step eight is repeated. Here, α is the redundancy coefficient, which can usually be taken as 1~1.05.

[0072] In one specific implementation, a composite material vertical tail on an aircraft suffered penetrating damage. Ultrasonic non-destructive testing was performed on the damage to identify the damaged area, and it was proposed to repair the damaged area by excavation and bonding.

[0073] The minimum safety margin for the damaged area is 0.2, and the key factor affecting the safety margin is shear performance. Therefore, an in-plane shear strength test was designed. The test specimen size and test plan were carried out in accordance with relevant ASTM standards to verify the safety of the repaired structure.

[0074] Metallographic testing of the damaged area repair test specimen showed an average porosity of less than 3%, while the metallographic testing of the damaged area calibration test specimen showed a porosity of less than 2%. In-plane shear strength tests were conducted, and the echo intensity loss Δdb1 of the damaged area repair test specimen relative to the damaged area calibration test specimen was calculated. Furthermore, the mechanical property reduction coefficient K, a key influencing factor, was 0.9.

[0075] The calculated repair safety margin RF = 0.08 indicates that the designed patch-and-adhesive repair scheme meets the requirements. Therefore, the patch-and-adhesive repair structure is adopted for the damaged area, and the repair patch structure is as follows: Figure 2 As shown, non-destructive testing is performed, and the echo intensity loss Δdb2 of the damaged area relative to the undamaged area is calculated. Since Δdb2 < Δdb1, the repair of the damaged area is completed.

[0076] The aforementioned method for repairing damage to composite material structures on aircraft, while ensuring manufacturing reliability, provides a method for judging the effectiveness of damage repair in service and maintenance scenarios by establishing comparative specimens of different dimensions and introducing a weakening coefficient. It also provides a multi-faceted, step-by-step judgment method for aspects such as the selection of repair materials and the control of repair processes, thus ensuring the reliability of the repair plan implementation.

[0077] The technical solution of this application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the scope of protection of this application is obviously not limited to these specific embodiments. Without departing from the principles of this application, those skilled in the art can make equivalent changes or substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will all fall within the scope of protection of this application.

Claims

1. A method for repairing damage to composite material structures on aircraft, characterized in that, include: Step 1: Determine the damaged area of ​​the structure, query the minimum safety margin RF0 of the damaged area, and the key factors affecting the safety margin, and then remove the damage. Step 2: Design a repair plan for the structure; Step 3: Prepare repair test pieces for the damaged area according to the repair plan, and conduct ultrasonic non-destructive testing, mechanical property testing of key influencing factors, and metallographic testing on the repair test pieces for the damaged area. Step 4: Fabricate calibration test pieces for the damaged area according to the materials and processes of the original structure, and conduct metallographic testing, ultrasonic non-destructive testing, and mechanical property tests on the calibration test pieces for the damaged area, as well as tests on key influencing factors. Step 5: Calculate the echo intensity loss Δdb1 of the repair test piece relative to the calibration test piece in the damaged area, and the key influencing factor, the mechanical property reduction coefficient K. Step 6: Using the minimum safety margin RF0 of the damaged area and the mechanical performance reduction coefficient K of the key influencing factor, calculate the repair safety margin RF of the damaged area and determine whether the repair plan meets the requirements. Step 7: Repair the damaged areas of the structure according to the repair plan; Step 8: Perform ultrasonic non-destructive testing on the structure, calculate the echo intensity loss Δdb2 of the damaged area relative to the undamaged area, compare it with Δdb1, and judge the repair effect on the damaged area.

2. The method for repairing damage to composite material structures on aircraft according to claim 1, characterized in that, In step one, ultrasonic non-destructive testing is performed on the structure, and the damaged area is determined by the echo intensity.

3. The method for repairing damage to composite material structures on aircraft according to claim 2, characterized in that, In step two, a repair plan for the structure is designed, including repair materials and repair processes.

4. The method for repairing damage to composite material structures on aircraft according to claim 3, characterized in that, In step three, multiple repair test pieces for damaged areas are made according to the repair plan. The echo intensity db1 of the ultrasonic non-destructive testing of the repair test pieces for damaged areas is taken as the average of the maximum and minimum values. The mechanical property test results of the key influencing factors of the repair test piece in the damaged area are recorded as F1.

5. The method for repairing damage to composite material structures on aircraft according to claim 4, characterized in that, In step three, metallographic examination is performed at the locations of the maximum and minimum echo intensities of the repair test piece in the damaged area.

6. The method for repairing damage to composite material structures on aircraft according to claim 5, characterized in that, In step three, if the metallographic test shows an average porosity greater than 3%, the repair plan is considered unacceptable and needs to be improved.

7. The method for repairing damage to composite material structures on aircraft according to claim 6, characterized in that, In step four, multiple damage area comparison test pieces are made according to the materials and processes of the original structure. The process allowance of each damage area comparison test piece is cut off, and metallographic testing is performed. The damage area comparison test piece with the smallest porosity obtained from the metallographic test is selected as the damage area calibration test piece.

8. The method for repairing damage to composite material structures on aircraft according to claim 7, characterized in that, In step four, the porosity obtained from metallographic testing of the calibrated test piece in the damaged area is less than 2%.

9. The method for repairing damage to composite material structures on aircraft according to claim 8, characterized in that, In step four, the echo intensity db2 of the ultrasonic nondestructive testing of the damaged area of ​​the calibrated test piece is taken as the average value of multiple locations; The mechanical property test results of the key influencing factors of the damaged area calibration test piece are recorded as F2.

10. The method for repairing damage to composite material structures on aircraft according to claim 9, characterized in that, In step five, the echo intensity loss Δdb1 of the repair test piece relative to the calibration test piece in the damaged area is calculated, specifically as follows: △db1 = db1 - db2.

11. The method for repairing damage to composite material structures on aircraft according to claim 10, characterized in that, In step five, the reduction coefficient K of the mechanical properties of the key influencing factors is calculated, specifically as follows: K = F1 / F2.

12. The method for repairing damage to composite material structures on aircraft according to claim 11, characterized in that, In step six, the repair safety margin RF of the damaged area is calculated, specifically as follows: RF = K × (RF0 + 1) - 1.

13. The method for repairing damage to composite material structures on aircraft according to claim 12, characterized in that, In step six, if the repair safety margin RF of the damaged area is less than 0, the repair plan is considered to be unacceptable and the repair plan will be improved.

14. The method for repairing damage to composite material structures on aircraft according to claim 13, characterized in that, In step eight, the echo intensity db3 of the ultrasonic nondestructive testing in the structural damage area is taken as the average value of multiple locations.

15. The method for repairing damage to composite material structures on aircraft according to claim 14, characterized in that, In step eight, the echo intensity db4 of the ultrasonic non-destructive testing of the undamaged area of ​​the structure is taken as the average value of multiple locations.

16. The method for repairing damage to composite material structures on aircraft according to claim 15, characterized in that, In step eight, the echo intensity loss Δdb2 of the damaged area relative to the undamaged area is calculated, specifically as follows: △db2 = db3 - db4.

17. The method for repairing damage to composite material structures on aircraft according to claim 16, characterized in that, In step eight, if △db2>α△db1, it is considered that the repair effect on the damaged area is not good, and the damaged area of ​​the structure is repaired again according to the repair plan, where α is the redundancy coefficient.

18. The method for repairing damage to composite material structures on aircraft according to claim 17, characterized in that, In step eight, the redundancy coefficient is taken as 1 to 1.05.

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