A method for repairing layered damage to aircraft panels

Abstract

This invention provides a layered damage repair method for aircraft panels. The method uses cracks as a reference. If the crack is near the inner side of the frame, the crack on the frame flange is completely removed along the direction parallel to the crack on the inner side of the frame. If the crack is near the outer side of the frame, the crack on the frame flange is completely removed along the direction parallel to the crack on the outer side of the frame. Non-destructive testing is performed on the frame area. An oxidizing solution is applied to the frame area. If the milled area is located on the inner side of the frame, a backing plate is manufactured, anodized with yellow pigment, and then coated with polyurethane primer. If the milled area is located on the outer side of the frame, a cover pressure ring is manufactured, milled, and replaced with the original cover pressure ring. If the milled area is located on the inner side of the frame, sealant is applied to the milled area. Sealant is applied to the contact surface between the backing plate and the milled area of ​​the frame, and then the backing plate is installed on the milled area of ​​the frame. This method can reduce the rapid response time for field repairs and improve aircraft safety and reliability.

Description

Technical Field

[0001] This invention belongs to the field of aircraft wing structure repair technology, and particularly relates to a method for repairing layered damage to aircraft panels. Background Technology

[0002] With the rapid development of digital design and manufacturing technologies in the aviation industry, integrated structural design has been widely adopted in aircraft structural design. Integral fuel tanks are crucial structural components of aircraft; damage and leakage during use can lead to serious consequences. Integral fuel tanks possess extremely high sealing performance while maintaining the aircraft's good shape and participating in the overall stress distribution. Therefore, effectively addressing the maintenance and repair of integral fuel tanks, and shortening maintenance and repair time, is of great significance for improving aircraft utilization, reducing flight hazards, and lowering maintenance and repair costs. The integral fuel tank's wall panel integrates the wing skin with stringers, vertical ribs, joints, and other possible structural components (such as fuel filler ports and inspection ports) into a single, integrally machined structure. This type of wall panel has a complex structure, with curved surfaces and requiring thick material blanks. Therefore, a process of first machining flat plates and then integrally bending is employed. However, bending is a highly complex elasto-plastic deformation process, and residual stress remains significant at the top and root of the ribs during the springback process. As aircraft operate in increasingly harsh environments (humid marine atmosphere, heavy salt spray, large waves, large temperature differences, etc.), materials are more sensitive to stress corrosion cracking under severe corrosive conditions, leading to stress corrosion cracks appearing on the inspection frame of the panel.

[0003] The integral wing panels of an aircraft have a lifespan equal to that of the aircraft. These panels typically feature elliptical maintenance access panels. When cracks appear in the access panel frame, traditional repair methods include cutting off the cracked section of the panel, creating a custom seal, or replacing the entire panel. However, cutting off the panel disrupts the main force transmission path of the integral wing panel; custom seals require new molds, which are costly; and replacing the panel poses significant safety risks: the sealant at the sealing ribs is difficult to remove; removing the cracked panel may cause deformation of the wing box, requiring new assembly fixtures to ensure proper functioning; and after replacing the panel, the two ends of the wing box section need to be re-machined, which may lead to out-of-tolerance hole diameters at the connection points, posing flight safety hazards. Summary of the Invention

[0004] To address the problems in existing technologies where cracks in aircraft door frames disrupt the main force transmission path by cutting off the panel, the high cost of specialized seals, and the significant risks associated with replacing the entire panel, this invention provides a method for repairing cracks in aircraft door frame panels. This method does not alter the structural force transmission path while meeting strength design requirements, thus solving the problems of interrupted force transmission due to panel cutting, high cost of specialized seals, and safety risks associated with replacing the entire panel. Furthermore, the method is simple, convenient, and quick, reducing field repair response time and improving aircraft safety and reliability. The technical solution is as follows:

[0005] In a first aspect, a method for repairing layered damage to aircraft panels is provided, the method comprising:

[0006] Step 1: Using the crack as a reference, if the crack is close to the inside of the frame, the crack on the flange of the frame is completely removed on the inside of the frame by milling along the direction parallel to the crack. After removal, the ends of the frame are rounded. If the crack is close to the outside of the frame, the crack on the flange of the frame is completely removed on the outside of the frame by milling along the direction parallel to the crack. After removal, the ends of the frame are rounded.

[0007] Step 2: Perform non-destructive testing on the removed frame area. If the test results are good and there is no internal or surface damage, proceed to step 3.

[0008] Step 3: Apply an oxidizing solution to the removed mouth frame area, and then spray multiple layers of polyurethane primer;

[0009] Step 4: If the milled part is located inside the mouth frame, manufacture a pad part. The shape of the pad part matches the milled shape inside the mouth frame. After yellow anodizing the pad part, spray a layer of polyurethane primer. If the milled part is located outside the mouth frame, manufacture a mouth cap ring part. Mill the mouth cap ring part to locally thicken the inside of the mouth cap ring part and match the milled shape outside the mouth frame. Replace the original mouth cap ring part with the milled mouth cap ring part. The repair work is completed in this case.

[0010] Step 5: If the milled part is located inside the frame, apply a layer of sealant to the milled part of the frame; proceed to step 6.

[0011] Step 6: Apply a layer of sealant to the contact surface between the pad part and the milled part of the frame to ensure the oil tank is sealed, and then install the pad part on the milled part of the frame.

[0012] Furthermore, the method also includes:

[0013] Step 7: Conduct a strength assessment on the repaired wall panel frame structure to determine whether it meets the strength requirements; if it meets the strength requirements, the wall panel frame structure repair is considered successful.

[0014] Optionally, in step 7,

[0015] When performing a strength assessment on the repaired wall panel frame structure, the frame is simplified as a beam element. The maximum axial force at the frame, F = σA, is calculated first. eq Then, based on the axial tensile and compressive equivalent cross-sectional area A of the repaired frame... eq The maximum axial stress at the repaired frame was calculated to be σ2 = F / A. eq Therefore, the safety margin MS = σ is calculated based on the maximum axial stress at the frame. 0.2 / σ2-1, if the safety margin MS is greater than 0, the wall panel frame structure repair is confirmed to be successful;

[0016] Where σ represents the tensile stress at the opening frame, and A eq σ represents the equivalent cross-sectional area of ​​the original frame under axial tension and compression. 0.2 This indicates the yield strength of the frame material.

[0017] Optionally, in step 1, the rounded corners at both ends of the removed frame are R10.

[0018] Optionally, in step 2, the non-destructive testing method is eddy current and ultrasonic testing.

[0019] Optionally, the polyurethane primer in steps 3 and 4 is S06-1010H polyurethane primer.

[0020] Optionally, the sealant used in steps 5 and 6 is XM-22 sealant.

[0021] Optionally, in step 6, the pad part is installed on the milled part of the frame using a double-sided countersunk annular groove sealing rivet.

[0022] Optionally, in step 4, the material of the pad plate part is the same as the material of the wall plate, and the material of the cap pressure ring part is the same as the material of the original cap pressure ring part.

[0023] The beneficial effects of this invention are at least as follows:

[0024] This invention provides a repair method that involves a custom-made pad made of the same material as the wall panel or a cover pressure ring similar to the original cover pressure ring. This method does not alter the structural force transmission path while meeting strength design requirements, solving the problems of interrupting the force transmission path by cutting off the wall panel, the high cost of custom-made seals, and the safety risks associated with replacing the entire wall panel. Furthermore, the method is simple, convenient, and quick, reducing field repair response time and improving aircraft safety and reliability. Attached Figure Description

[0025] Figure 1 This is a partial schematic diagram of the inspection frame for the upper wing panel;

[0026] Figure 2A typical schematic diagram of the fit between the frame and cover of the upper wing panel;

[0027] Figure 3 To strengthen the minimum cross-sectional dimensions of the frame;

[0028] Figure 4 The crack length and depth;

[0029] Figure 5 The location of the crack and the condition of the polishing;

[0030] Figure 6 For the milling area of ​​the opening frame;

[0031] Figure 7 This is a schematic diagram of the pad component;

[0032] Figure 8 This is a schematic diagram showing the connection between the wall panel and the pad plate.

[0033] Figure 9 Schematic diagram of the cap pressure ring component;

[0034] Figure 10 This is a flowchart of the aircraft panel layered damage repair method of the present invention. Detailed Implementation

[0035] The present invention will now be described in further detail with reference to specific embodiments and accompanying drawings.

[0036] See Figure 10 An embodiment of the present invention provides a method for repairing layered damage to aircraft panels, the method comprising the following steps:

[0037] Step 1: Using the crack as a reference, if the crack is close to the inside of the frame, the crack on the flange of the frame is completely removed on the inside of the frame by milling along the direction parallel to the crack. After removal, the ends of the frame are rounded. If the crack is close to the outside of the frame, the crack on the flange of the frame is completely removed on the outside of the frame by milling along the direction parallel to the crack. After removal, the ends of the frame are rounded.

[0038] Step 2: Perform non-destructive testing on the removed frame area. If the test results are good and there is no internal or surface damage, proceed to step 3.

[0039] Step 3: Apply an oxidizing solution to the removed mouth frame area, and then spray multiple layers of polyurethane primer;

[0040] Step 4: If the milled part is located inside the mouth frame, manufacture a pad part. The shape of the pad part matches the milled shape inside the mouth frame. After yellow anodizing the pad part, spray a layer of polyurethane primer. If the milled part is located outside the mouth frame, manufacture a mouth cap ring part. Mill the mouth cap ring part to locally thicken the inside of the mouth cap ring part and match the milled shape outside the mouth frame. Replace the original mouth cap ring part with the milled mouth cap ring part. The repair work is completed in this case.

[0041] Step 5: If the milled part is located inside the frame, apply a layer of sealant to the milled part of the frame; proceed to step 6.

[0042] Step 6: Apply a layer of sealant to the contact surface between the pad part and the milled part of the frame to ensure the oil tank is sealed, and then install the pad part on the milled part of the frame.

[0043] Example 1:

[0044] Crack description: Fuel tank inspection port frame (material: 1973T2) on the upper right wing panel, ribs 4-5. Figure 1 , Figure 2 Its typical circumferential cross-section is shown below. Figure 3 A significant circumferential crack was found on the side end face. After eddy current and ultrasonic testing, the preliminary assessment of the crack failure is as follows: (1) The crack length is approximately 100 mm. Figure 4 The maximum depth is about 15mm, and the depth is uniform except at both ends. The minimum distance between the crack and the end face of the stringer is 18mm. (2) The crack is about 9mm from the upper surface of the frame and about 6mm from the lower surface of the frame. The height of the crack from the upper and lower surfaces on the side end face of the frame remains basically constant along the circumference of the frame. See Figure 5 .

[0045] Taking the aforementioned crack as an example, the method for repairing layered damage to aircraft panels provided by this invention may specifically include the following steps:

[0046] 1) Based on the crack location, using the crack as a reference, completely mill away the crack in the flange of the mouth frame along the crack direction on the inner side of the mouth frame. Mill at a depth of 19mm, while ensuring a smooth transition of R4 at the sharp corners; ensure the crack is completely removed along the crack length direction, with a smooth transition of R10 at both ends. See [reference needed]. Figure 6 ;

[0047] 2) After the flange of the mouth frame is milled, non-destructive testing is performed on the milled part. If there is no internal or surface damage, the procedure is continued according to this plan; if there is damage, the procedure in 1) must be repeated.

[0048] 3) After inspection, apply an oxidizing solution to the milled area, and then spray multiple layers of S06-1010H polyurethane primer.

[0049] 4) Press Figure 7 A special pad part is made. The special pad material is 1973T2, and the shape is determined according to the spanning milling dimensions. After the pad is anodized in yellow, it is then coated with S06-1010H polyurethane primer.

[0050] 5) Apply a layer of XM-22 sealant to the milled area of ​​the mouth frame;

[0051] 6) Before installing the pad and wall panel frame, apply a layer of XM-22 sealant to the contact surface according to the relevant technical conditions (manufacturing technical conditions of the central wing and the integral fuel tank of the middle and outer wings, etc.) to ensure the fuel tank is sealed.

[0052] 7) Specially made pad parts according to Figure 8 The connection to the frame is achieved using five double-sided countersunk annular groove sealing rivets. The two end rivets are 9mm apart along the crack length, with four rivets spaced equidistantly in the middle. The other five rivets are 7mm apart along the crack depth. After the shim is connected, excess sealant at the frame is removed. When countersunk the rivets at the connection surface between the frame and the metal sealing ring, the countersunk angle is 90°, the countersunk diameter D = 5mm, and the depth is less than 1.7mm. To ensure proper installation of the sealing ring, the double-sided countersunk rivet connection should be smooth and flat, without burrs, and flush with the original frame. When the countersunk diameter is φ5, the portion entering the R-zone of the frame should transition smoothly.

[0053] 8) Strength Analysis

[0054] When performing a strength assessment on the repaired wall panel frame structure, the frame is simplified as a beam element. The maximum axial force at the frame, F = σA, is calculated first. eq Then, based on the axial tensile and compressive equivalent cross-sectional area A of the repaired frame... eq The maximum axial stress at the repaired frame was calculated to be σ2 = F / A. eq Therefore, the safety margin MS = σ is calculated based on the maximum axial stress at the frame. 0.2 / σ2-1, if the safety margin MS is greater than 0, the wall panel frame structure repair is confirmed to be successful;

[0055] Where σ represents the tensile stress at the opening frame, and A eq σ represents the equivalent cross-sectional area of ​​the original frame under axial tension and compression. 0.2 This indicates the yield strength of the frame material.

[0056] a) Qualitative analysis:

[0057] In the impact analysis, the outer wing No. 2 upper panel is a one-piece machined part, and the material is 1973T2.

[0058] (σ 0.2 =441MPa), the specific structure is as follows Figure 1 As shown.

[0059] Depend on Figure 1As can be seen, the two long stringers of the wing break off at both sides of the inspection frame. To transfer the loads of the stringers and the wall panels, the inspection frame is locally reinforced. The minimum cross-sectional dimension of the reinforced inspection frame is as follows: Figure 3 As shown, the reinforced inspection frame mainly bears axial tensile and compressive forces.

[0060] Depend on Figure 5 It can be seen that the milled area is located near the center line of the opening frame. This area mainly serves to fix the inspection cover and mainly bears the axial tensile and compressive stress of the opening frame.

[0061] After milling away the cracked area, the cross-sectional area of ​​the frame decreases, and the axial tensile and compressive stress of the frame increases.

[0062] b) Strength assessment:

[0063] Specifically, the strength assessment includes the following steps:

[0064] 1. Calculate the maximum axial force at the opening frame:

[0065] According to the document "Finite Element Model and Stress Analysis of Outer Wing of a Certain Type of Aircraft", the maximum axial tensile stress at the finite element full-machine solution of the frame is σ = 137.93 MPa and the maximum axial compressive stress is σ = 254.72 MPa.

[0066] Depend on Figure 3 It can be seen that the minimum cross-sectional area of ​​the reinforcing frame is:

[0067] A=A1+A2+A3=(33-4.3-15)×21+15×(19+21)+4.3×(19-14+21)=999.5mm 2 .

[0068] Centroid of section:

[0069]

[0070] The distance from the bottom of the cross-section to the chord line is h0 = 74.37 mm. Therefore, the equivalent cross-sectional area is:

[0071] A eq =A×((18.15+74.37) / (33+74.37)) 2 =731.6mm 2 .

[0072] Maximum axial force at the frame: F = σA eq =254.72×731.6=186353.15N.

[0073] 2. Calculate the equivalent cross-sectional area after repair:

[0074] Depend on Figure 5 The minimum cross-sectional area of ​​the frame after milling is known to be:

[0075] A'=A1'+A2'+A3'=21×(33-4.3-15+6)+(19+21)×(15-6)+(21+19-14)

[0076] ×4.3=885.50mm 2 .

[0077] Centroid of section:

[0078]

[0079] Equivalent cross-sectional area:

[0080] A eq '=A'×((18.34+74.37) / (33+74.37)) 2 =660.20mm 2 .

[0081] Calculations show that the minimum cross-sectional area of ​​the frame is reduced by 11.41%, and the minimum equivalent cross-sectional area of ​​the frame is reduced by 9.76%.

[0082] 3. Calculate the maximum axial force and safety margin (MS) after repair.

[0083] σ²=F / A eq '=186353.15 / 660.20=282.27MPa.

[0084] MS = σ 0.2 / σ2-1=441 / 282.27-1=0.56.

[0085] The wall panel frame structure has been successfully repaired, meeting the strength design requirements.

[0086] If the crack is near the outer edge of the mouth frame, then it can be handled according to... Figure 9 The specially made cap pressure ring part was repaired.

[0087] The above description merely illustrates embodiments of the present invention and is quite specific and detailed; however, it should not be construed as limiting the scope of the patent. It should be noted that those skilled in the art can make various modifications and improvements without departing from the concept of the present invention, and these all fall within the scope of protection of the present invention. Furthermore, any parts of the present invention not described in detail are conventional techniques.

Claims

1. A method for repairing layered damage to aircraft panels, characterized in that, The method includes: Step 1: Using the crack as a reference, if the crack is close to the inside of the frame, the crack on the flange of the frame is completely removed on the inside of the frame by milling along the direction parallel to the crack. After removal, the ends of the frame are rounded. If the crack is close to the outside of the frame, the crack on the flange of the frame is completely removed on the outside of the frame by milling along the direction parallel to the crack. After removal, the ends of the frame are rounded. Step 2: Perform non-destructive testing on the removed frame area. If the test results are good and there is no internal or surface damage, proceed to step 3. Step 3: Apply an oxidizing solution to the removed mouth frame area, and then spray multiple layers of polyurethane primer; Step 4: If the milled part is located inside the mouth frame, manufacture a pad part. The shape of the pad part matches the milled shape inside the mouth frame. After yellow anodizing the pad part, spray a layer of polyurethane primer. If the milled part is located outside the mouth frame, manufacture a mouth cap ring part. Mill the mouth cap ring part to locally thicken the inside of the mouth cap ring part and match the milled shape outside the mouth frame. Replace the original mouth cap ring part with the milled mouth cap ring part. The repair work is completed in this case. Step 5: If the milled area is located inside the frame, apply a layer of sealant to the milled area of ​​the frame; proceed to step 6. Step 6: Apply a layer of sealant to the contact surface between the pad part and the milled part of the frame to ensure the oil tank is sealed, and then install the pad part on the milled part of the frame.

2. The method according to claim 1, characterized in that, The method further includes: Step 7: Conduct a strength assessment on the repaired wall panel frame structure to determine whether it meets the strength requirements; if it meets the strength requirements, the wall panel frame structure repair is considered successful.

3. The method according to claim 2, characterized in that, In step 7, When performing a strength assessment on the repaired wall panel frame structure, the frame is simplified as a beam element. The maximum axial force at the frame, F = σA, is calculated first. eq Then, based on the axial tensile and compressive equivalent cross-sectional area A of the repaired frame... eq The maximum axial stress at the repaired frame was calculated to be σ2 = F / A. eq Therefore, the safety margin MS = σ is calculated based on the maximum axial stress at the frame. 0.2 / σ2-1, if the safety margin MS is greater than 0, the wall panel frame structure repair is confirmed to be successful; Where σ represents the tensile stress at the opening frame, and A eq σ represents the equivalent cross-sectional area of ​​the original frame under axial tension and compression. 0.2 This indicates the yield strength of the frame material.

4. The method according to claim 1, characterized in that, In step 1, the rounded corners at both ends of the rear frame are removed to R10.

5. The method according to claim 1, characterized in that, In step 2, the non-destructive testing methods are eddy current and ultrasonic testing.

6. The method according to claim 1, characterized in that, The polyurethane primer used in steps 3 and 4 is S06-1010H polyurethane primer.

7. The method according to claim 1, characterized in that, The sealant used in steps 5 and 6 is XM-22 sealant.

8. The method according to claim 1, characterized in that, In step 6, the pad plate part is installed on the milled part of the frame using double-sided countersunk annular groove sealing rivets.

9. The method according to claim 1, characterized in that, In step 4, the material of the pad plate part is the same as that of the wall plate, and the material of the cap pressure ring part is the same as that of the original cap pressure ring part.

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