Aircraft skin longitudinal joint structure strength analysis method

Abstract

This invention designs a method for strength analysis of longitudinal connection structures in aircraft skin, belonging to the field of airframe structural strength design. Through finite element modeling and experimental analysis, it determines the load-sharing factor of fasteners and the influence of secondary bending on the load-bearing capacity of the connection area. This invention significantly improves the accuracy of strength calculations for longitudinal connection structures in aircraft skin, providing support for strength analysis of aircraft products. This invention has wide applications in structural design in aviation, aerospace, and automotive industries; it can improve structural utilization efficiency, designing optimal structural dimensions while meeting structural strength requirements; it helps reduce unnecessary conservative designs, greatly shortening development time and reducing development costs.

Description

Technical Field

[0001] This invention belongs to the field of airframe structural strength design, and particularly relates to a method for strength analysis of longitudinal connection structures of aircraft skin. Background Technology

[0002] The longitudinal skin connection structure of the fuselage is a critical connection area of ​​an aircraft, with complex load transfer patterns, making effective strength analysis a consistently challenging task for engineers. Currently in service worldwide, the longitudinal skin connections of aircraft fuselages generally fall into two structural forms: overlapping skin and butt-jointed skin. Each connection configuration has its own advantages and disadvantages, and the appropriate choice should be made based on the specific circumstances. Figure 1 There are two structural forms for longitudinal skin connections on the fuselage. Form A is the longitudinal lap joint configuration, which uses fewer parts and fasteners and is simpler to assemble than the butt joint configuration. Therefore, it is widely used in areas other than the forward fuselage and mid-fuselage upper panels. Furthermore, the lap joint is generally an upper-overlap configuration, with the upper skin on the outside and the lower skin on the inside, for rain protection. Form B is the longitudinal skin butt joint configuration, typically used in the forward fuselage and mid-fuselage upper panels, to account for the impact of the joint on the fuselage's aerodynamic loads.

[0003] Longitudinal connections in wall panels are critical connection areas, transmitting significant loads. Mechanical fastener connections are typically used to ensure connection performance. Common failure modes in longitudinal connections include cracking of the skin at bosses and nail holes at the joints, and fastener failure preventing load transfer between the connected components. Currently, there is no unified and effective calculation method for the failure modes and failure loads of longitudinal connection structures. Most methods simply distribute the external load evenly across each fastener, only checking the fastener strength under shear loads, without fully considering the load distribution factor of each fastener or the impact of secondary bending caused by lap or butt joint eccentricity on the loads on the fasteners and skin. Summary of the Invention

[0004] To address the shortcomings of existing technologies, this invention presents a method for strength analysis of longitudinal connection structures in aircraft skin.

[0005] A method for strength analysis of longitudinal connection structures of aircraft skin, comprising the following steps:

[0006] Step 1: Using an Excel tool, extract the internal loads of the longitudinal connection structure of the skin from the calculation results of the overall finite element model of the aircraft; where F x1 F y1 q xy1 and F x2 F y2 q xy2 These represent the force flows of the skin elements on the upper and lower sides of the connection area in the local coordinate system, P. G1P G2 These are the end loads at both ends of the stringer;

[0007] Step 2: Determine the allowable strength values ​​for aircraft fasteners and skin; P tu and P su These are the ultimate tensile strength and ultimate shear strength of fasteners, σ tu and σ cy These are the tensile ultimate strength and compressive yield strength of the skin material, respectively.

[0008] Step 3: Calculate the load distribution between the fasteners and the connecting plate using the constructed equilibrium equations and deformation compatibility equations. If there are n rows of fasteners, there is one equilibrium equation and n-1 deformation compatibility equations. Use a single-shear connector with two rows of fasteners to illustrate the parameters in the equilibrium and compatibility equations. The first row of fasteners is subjected to a load V1, which is transferred to the connecting plate A, i.e., the upper skin, i.e., P. A =V1; simultaneously, the remaining load (P-V1) is transferred to the connecting strip plate B, i.e., the lower skin, i.e., P. B =P-V1, and is borne by the last row of fasteners, so V2 = P-V1; therefore, the total load is:

[0009] P = V1 + V2 = P A +P B (1)

[0010] Based on deformation compatibility, the total deformation of fastener 1 and connecting plate A is equal to the total deformation of fastener 2 and connecting plate B:

[0011]

[0012] V1C1+δ A =V2C2+δ B (3)

[0013] The load P of the connecting plate can then be calculated. A P B and the loads V1 and V2 of the fasteners;

[0014] In the formula, δ1 and δ2 represent the deformation of the fastener, δ A δ B For the deformation of the connecting strip, L is the length of the connecting strip, A is the cross-sectional area of ​​the connecting strip, E is the elastic modulus of the connecting strip, and C is the flexibility constant of the fastener. Refer to Formula 4:

[0015]

[0016] In the formula, t spA t spB Where is the thickness of the connecting plate unit; D is the diameter of the fastener; B1 and B2 are constants.

[0017] Step 4: Calculate the load distribution factor for the fasteners and connecting strips;

[0018] It is the load-sharing factor of the fastener in the nth row;

[0019] and These are the load-sharing factors for connecting plates A and B, respectively;

[0020] Step 5: Establish a detailed finite element model of the longitudinal connection structure of the skin in the finite element simulation software, extract the shear load of the fastener and the tensile load caused by secondary bending, and calculate the load distribution factor f of the shear and tensile loads of the fastener. s f t ;

[0021] f s = Fastener shear load / (P·f) n )

[0022] f t = Fastener tensile load / (P·f) n )

[0023] Step 6: Calculate the tensile load on the fastener based on Step 3;

[0024]

[0025] In the formula, t spA , t spB These represent the thicknesses of the connecting strip unit and D represents the diameter of the fastener.

[0026] Step 7: Calculate the shear load on the fastener based on Step 3:

[0027]

[0028] P y =d pitch ·F y (7)

[0029] V xy =d pitch ·q xy (8)

[0030] In the formula, n is the number of rows of fasteners in the connection area, and d pitch It is the longitudinal fastener spacing, F y and q xy These are the circumferential force flow and shear flow borne by the skin, respectively;

[0031] Step 8: Calculate the safety margin of the fasteners:

[0032]

[0033] In the formula: It is the ratio of the actual tensile load on the fastener to the allowable tensile value; It is the ratio of the actual shear load on the fastener to the allowable shear value; P tu It is the ultimate tensile strength of the fastener; P su It is the ultimate shear strength of the fastener;

[0034] Step 9: Establish a detailed finite element model of the longitudinal connection structure of the skin, and calculate the secondary bending coefficient f of the connecting strip plate. BM :

[0035]

[0036] σ axial and σ bend In the finite element model, the axial stress and bending stress at the same location on the skin are represented.

[0037] Step 10: Calculate the safety margins for the skin boss locations and connecting strip locations; safety margins under maximum and minimum principal stresses:

[0038]

[0039]

[0040] Where: σ tu and σ cy These are the tensile ultimate strength and compressive yield strength of the material, σ max and σ min These are the maximum principal stress and the minimum principal stress, respectively.

[0041] Stress calculation at the skin boss:

[0042]

[0043]

[0044] In the formula, t p =min(t) pA ,t pB ),

[0045] A ska =(t spA +t spB )×e B +t spA ×d A +t pA ×min(W pA W a / 2-d A-e B )+t A ×max(W a / 2-d A -e B -W pA ,0)

[0046] A skb =(t spA +t spB )×(W sp -e B )+t spB ×d B +t pB ×min(W pB W b / 2+e B -W sp -d B )+t B ×max(W b / 2+e B -W sp -d B -W pB ,0)

[0047] Among them, t spA and t spB The thickness e of the connecting strips A and B A and e B : The distance between the nail edges connecting plates A and B; d A and d B : The distance between the edges of connecting plates A and B; t pA and t pB : Thickness of skin bosses A and B; W pA and W pB : Width of skin bosses A and B; W a and W b : Spacing between the stringers of skins A and B; t A and t B : Thickness of skin A and B; W sp Width of the connection area;

[0048] Maximum / minimum principal stress at the skin boss:

[0049]

[0050] Consider the maximum and minimum principal stresses under different stress combinations:

[0051] σ1=σ_a x , σ2=σ_a y , σ3=τ_a xy

[0052] σ1=σ_b x , σ2=σ_b y , σ3=τ_b xy

[0053] Stress calculation at skin joints:

[0054]

[0055]

[0056] In the formula, t sp =min(t) spA ,t spB )

[0057] Maximum / minimum principal stress at skin joints:

[0058]

[0059] Consider the maximum and minimum principal stresses under different stress combinations:

[0060] σ1=σ_a x , σ2=σ_a y , σ3=τ_a xy

[0061] σ1=σ_b x , σ2=σ_b y , σ3=τ_b xy

[0062] Step 11: Calculate the safety margin at the nail hole positions of the connecting strip; the maximum bending stress caused by the secondary bending occurs at the first row of fasteners in the overlap area. Read the stress value σ under axial load at the nail hole position in the detailed finite element model. The maximum stress at the nail hole position is then:

[0063]

[0064]

[0065] In the formula, σ bru This refers to the material's compressive strength.

[0066] Beneficial technical effects of the present invention:

[0067] Based on research into the strength analysis methods of longitudinal connection structures in aircraft aft fuselages, this invention proposes a structural strength analysis method applicable to fuselage panel connection areas. This method determines the load-sharing factor of fasteners and the influence of secondary bending on the load-bearing capacity of the connection area structure through finite element modeling and experimental analysis. This invention significantly improves the accuracy of strength calculations for longitudinal connection structures in the skin, providing support for strength analysis of aircraft products.

[0068] The invention has wide applications in structural design in aviation, aerospace, and automotive industries; it can improve the utilization efficiency of structures and design the optimal structural dimensions while meeting structural strength requirements; it helps to reduce unnecessary conservative designs, greatly shortens research and development time and reduces research and development costs; it is conducive to exploring the structural weight reduction capacity and reducing transportation costs; the invention has been applied to the design of the CR929 aircraft, verifying its analytical accuracy and reliability, and has broad prospects for practical engineering applications. Attached Figure Description

[0069] Figure 1 This embodiment of the invention is a schematic diagram defining the dimensions of the longitudinal connection structure of the skin.

[0070] Figure 2 This embodiment of the invention is a schematic diagram of the internal load of the longitudinal connection structure of the skin.

[0071] Figure 3 This invention provides a schematic diagram of the longitudinal connection structure of the skin under load and deformation.

[0072] Figure 4 This embodiment of the invention is a schematic diagram of the location of maximum stress caused by a secondary bending.

[0073] Figure 5 The embodiment of the present invention is a deformation cloud diagram of the longitudinal connection structure of the skin.

[0074] Figure 6 The embodiments of the present invention address the tensile and shear loads on fasteners of longitudinal connection structures for skin.

[0075] Figure 7 The embodiment of the present invention is the stress value near the nail hole of the longitudinal connection structure of the skin.

[0076] Figure 8 This invention provides a diagram of a test fixture for a longitudinal connection structure of a skin. Detailed Implementation

[0077] The present invention will be further described below with reference to the accompanying drawings and embodiments;

[0078] This invention provides an analytical method applicable to longitudinal overlapping panel structures of aircraft skin, offering a theoretical basis for analysis results. It provides top-level design guidelines and analytical methods for longitudinal connection structures of civil aircraft panels, addressing the shortcomings in the accuracy of strength calculations for longitudinal connection structures of skin in China. This invention offers a more accurate and versatile calculation method, effectively solving the impact of fastener load distribution and secondary bending on calculation accuracy.

[0079] A method for strength analysis of longitudinal connection structures of aircraft skin, with the attached diagram of the test fixture for the longitudinal connection structure of the skin. Figure 8As shown; the steps are as follows:

[0080] Step 1: Using an Excel tool, extract the internal loads of the longitudinal connection structure of the skin from the calculation results of the overall finite element model of the aircraft; where F x1 F y1 q xy1 and F x2 F y2 q xy2 These represent the force flows of the skin elements on the upper and lower sides of the connection area in the local coordinate system, P. G1 P G2 The end loads at both ends of the stringer; a schematic diagram of the loads within the longitudinal connection structure of the skin is attached. Figure 2 As shown; a schematic diagram of the longitudinal connection structure of the skin under load and deformation is attached. Figure 3 As shown; the deformation contour diagram of the longitudinal connection structure of the skin is attached. Figure 5 As shown;

[0081] Step 2: Determine the allowable strength values ​​for aircraft fasteners and skin; P tu and P su These are the ultimate tensile strength and ultimate shear strength of fasteners, σ tu and σ cy These are the tensile ultimate strength and compressive yield strength of the skin material; the tensile and shear loads of the fasteners for the longitudinal connection structure of the skin are shown in the attached figure. Figure 6 As shown;

[0082] Step 3: Calculate the load distribution between the fasteners and the connecting plate using the constructed equilibrium equations and deformation compatibility equations. If there are n rows of fasteners, there is one equilibrium equation and n-1 deformation compatibility equations. Use a single-shear connector with two rows of fasteners to illustrate the parameters in the equilibrium and compatibility equations. The first row of fasteners is subjected to a load V1, which is transferred to the connecting plate A, i.e., the upper skin, i.e., P. A =V1; simultaneously, the remaining load (P-V1) is transferred to the connecting strip plate B, i.e., the lower skin, i.e., P. B =P-V1, and is borne by the last row of fasteners, so V2 = P-V1; therefore, the total load is:

[0083] P = V1 + V2 = P A +P B (1)

[0084] Based on deformation compatibility, the total deformation of fastener 1 and connecting plate A is equal to the total deformation of fastener 2 and connecting plate B:

[0085]

[0086] V1C1+δ A =V2C2+δB (3)

[0087] The load P of the connecting plate can then be calculated. A P B and the loads V1 and V2 of the fasteners;

[0088] In the formula, δ1 and δ2 represent the deformation of the fastener, δ A δ B For the deformation of the connecting strip, L is the length of the connecting strip, A is the cross-sectional area of ​​the connecting strip, E is the elastic modulus of the connecting strip, and C is the flexibility constant of the fastener. Refer to Formula 4:

[0089]

[0090] In the formula, t spA t spB For the thickness of the connecting plate unit, see attached. Figure 1 As shown; D is the diameter of the fastener, and B1 and B2 are constants; see Table 1.

[0091] Table 1 Fastener flexibility constants;

[0092] fastener B1 B2 aluminum alloy 5 0.8 steel 1.67 0.86 Titanium alloy 4 0.82

[0093] Step 4: Calculate the load distribution factor for the fasteners and connecting strips;

[0094] It is the load-sharing factor of the fastener in the nth row;

[0095] and These are the load-sharing factors for connecting plates A and B, respectively;

[0096] Step 5: Establish a detailed finite element model of the longitudinal connection structure of the skin in the finite element simulation software, extract the shear load of the fastener and the tensile load caused by secondary bending, and calculate the load distribution factor f of the shear and tensile loads of the fastener. s f t The attached diagram shows the location of the maximum stress caused by the secondary bending. Figure 4 As shown;

[0097] f s = Fastener shear load / (P·f) n )

[0098] f t = Fastener tensile load / (P·f) n )

[0099] Step 6: Calculate the tensile load on the fastener based on Step 3;

[0100]

[0101] In the formula, t spA , t spB These represent the thicknesses of the connecting strip unit and D represents the diameter of the fastener.

[0102] Step 7: Calculate the shear load on the fastener based on Step 3:

[0103]

[0104] P y =d pitch ·F y (7)

[0105] V xy =d pitch ·q xy (8)

[0106] In the formula, n is the number of rows of fasteners in the connection area, and d pitch It is the longitudinal fastener spacing, F y and q xy These are the circumferential force flow and shear flow borne by the skin, respectively;

[0107] Step 8: Calculate the safety margin of the fasteners:

[0108]

[0109] In the formula: It is the ratio of the actual tensile load on the fastener to the allowable tensile value; It is the ratio of the actual shear load on the fastener to the allowable shear value; P tu It is the ultimate tensile strength of the fastener; P su It is the ultimate shear strength of the fastener;

[0110] Step 9: Establish a detailed finite element model of the longitudinal connection structure of the skin, and calculate the secondary bending coefficient f of the connecting strip plate. BM :

[0111]

[0112] σ axial and σ bend In the finite element model, the axial stress and bending stress at the same location on the skin are represented.

[0113] Step 10: Calculate the safety margins for the skin boss locations and connecting strip locations; refer to Formulas 11 and 12 for the safety margins under the maximum and minimum principal stresses.

[0114]

[0115]

[0116] Where: σ tu and σ cy These are the tensile ultimate strength and compressive yield strength of the material, σ max and σ min These are the maximum principal stress and the minimum principal stress, respectively.

[0117] Stress calculation at the skin boss:

[0118]

[0119]

[0120] In the formula, t p =min(t) pA ,t pB ),

[0121] A ska =(t spA +t spB )×e B +t spA ×d A +t pA ×min(W pA W a / 2-d A -e B )+t A ×max(W a / 2-d A -e B -W pA ,0)

[0122] A skb =(t spA +t spB )×(W sp -e B )+t spB ×d B +t pB ×min(W pB W b / 2+e B -W sp -d B )+t B ×max(W b / 2+e B -W sp -d B -W pB ,0)

[0123] Among them, t spA and t spB The thickness e of the connecting strips A and BA and e B : The distance between the nail edges connecting plates A and B; d A and d B : The distance between the edges of connecting plates A and B; t pA and t pB : Thickness of skin bosses A and B; W pA and W pB : Width of skin bosses A and B; W a and W b : Spacing between the stringers of skins A and B; t A and t B : Thickness of skin A and B; W sp Width of the connection area;

[0124] Maximum / minimum principal stress at the skin boss:

[0125]

[0126] Consider the maximum and minimum principal stresses under different stress combinations:

[0127] σ1=σ_a x , σ2=σ_a y , σ3=τ_a xy

[0128] σ1=σ_b x , σ2=σ_b y , σ3=τ_b xy

[0129] Stress calculation at skin joints:

[0130]

[0131]

[0132] In the formula, t sp =min(t) spA ,t spB )

[0133] Maximum / minimum principal stress at skin joints:

[0134]

[0135] Consider the maximum and minimum principal stresses under different stress combinations:

[0136] σ1=σ_a x , σ2=σ_a y , σ3=τ_a xy

[0137] σ1=σ_b x, σ2=σ_b y , σ3=τ_b xy

[0138] Step 11: Calculate the safety margin at the nail hole positions of the connecting strip; the maximum bending stress caused by the secondary bending occurs at the first row of fasteners in the overlap area. Read the stress value σ under axial load at the nail hole position in the detailed finite element model. The maximum stress at the nail hole position is then:

[0139]

[0140]

[0141] In the formula, σ bru This represents the material's compressive strength. The stress values ​​near the nail holes in the longitudinal connection structure of the skin are shown in the attached figure. Figure 7 As shown;

[0142] The method of the present invention has the following functions:

[0143] Determine the constitutive equations for the strength analysis of the longitudinal connection structure of civil aircraft skin;

[0144] The internal loads of the longitudinal connection structure of the civil aircraft skin were calculated using a developed tool.

[0145] Determine the load distribution ratio of the fasteners in the longitudinal connection structure of the civil aircraft skin;

[0146] Determine the load distribution ratio of the connecting strip plates of the longitudinal connection structure of civil aircraft skin;

[0147] Determine the tensile load on fasteners caused by secondary bending in the longitudinal connection structure of civil aircraft skin;

[0148] Determine the effect of secondary bending on skin stress in the longitudinal connection structure of civil aircraft skin;

[0149] To determine the effect of secondary bending on the stress at the nail hole location in the longitudinal connection structure of civil aircraft skin.

Claims

1. A method for strength analysis of longitudinal connection structures of aircraft skin, characterized in that, The steps are as follows: Step 1: Using an Excel tool, extract the internal loads of the longitudinal connection structure of the skin from the calculation results of the overall finite element model of the aircraft; among which, , , and , , These represent the force flows of the skin elements on the upper and lower sides of the connection area in the local coordinate system. , These are the end loads at both ends of the stringer; Step 2: Determine the allowable strength values ​​for aircraft fasteners and skin; and These are the ultimate tensile strength and ultimate shear strength of fasteners, respectively. and These are the tensile ultimate strength and compressive yield strength of the skin material, respectively. Step 3: Calculate the load distribution between the fasteners and the connecting plate using the constructed equilibrium equations and deformation compatibility equations; with n rows of fasteners, there is one equilibrium equation and n-1 deformation compatibility equations; use a single shear connector with two rows of fasteners to illustrate the parameters in the equilibrium and compatibility equations, where the first row of fasteners is subjected to a load. This is transmitted to the connecting plate A, i.e., the upper skin. Meanwhile, the remaining load This is transmitted to the connecting plate B, i.e., the lower skin. And it is borne by the last row of fasteners, so P represents the total load; Step 4: Calculate the load distribution factor for the fasteners and connecting strips; It is the load-sharing factor of the fastener in the nth row; and These are the load-sharing factors for connecting plates A and B, respectively; Step 5: Establish a detailed finite element model of the longitudinal connection structure of the skin in the finite element simulation software, extract the shear load of the fastener and the tensile load caused by secondary bending, and calculate the load distribution factor of the shear and tensile loads of the fastener. , ; = Fastener shear load / ( ); = Fastener tensile load / ( ); Step 6: Calculate the tensile load on the fastener based on Step 3; The tensile load of the fastener is specifically: (5); In the formula, , These represent the thicknesses of the connecting strip unit and D represents the diameter of the fastener. Step 7: Calculate the shear load on the fastener based on Step 3, specifically: （6）； （7）； （8）； In the formula, n is the number of rows of fasteners in the connection area. It refers to the longitudinal fastener spacing. and These are the circumferential force flow and shear flow borne by the skin, respectively; Step 8: Calculate the safety margin of the fasteners; Step 9: Establish a detailed finite element model of the longitudinal connection structure of the skin, and calculate the secondary bending coefficient of the connecting strip plate. Specifically: （10）； and In the finite element model, the axial stress and bending stress at the same location on the skin are represented. Step 10: Calculate the safety margins for the positions of the skin boss and the connecting strip, specifically: Safety margins at maximum and minimum principal stresses: （11）； （12）； In the formula: and These are the tensile strength and compressive yield strength of the material, respectively. and These are the maximum principal stress and the minimum principal stress, respectively. Stress calculation at the skin boss: （13）； （14）； In the formula, , ； ； in, and Thickness of connecting strips A and B and : The distance between the nail edges connecting plates A and B; and : The distance between the edges of connecting plates A and B; and The thickness of skin bosses A and B; and : Width of skin bosses A and B; and : Spacing between the stringers of skins A and B; and The thickness of skins A and B; Width of the connection area; Maximum / minimum principal stress at the skin boss: （15）； Consider the maximum and minimum principal stresses under different stress combinations: ， ， ； ， ， ； Stress calculation at skin joints: （16）； （17）； In the formula, ; Maximum / minimum principal stress at skin joints: （18）； Consider the maximum and minimum principal stresses under different stress combinations: ， ， ； ， ， ； Step 11: Calculate the safety margin at the location of the nail holes in the connecting strip; the maximum bending stress caused by the secondary bending occurs at the location of the first row of fasteners in the overlap area. Read the stress value under axial load at the location of the nail hole in the detailed finite element model. The maximum stress at the nail hole location is: （19）； （20）； In the formula, This refers to the material's compressive strength.

2. The method for strength analysis of longitudinal connection structure of aircraft skin according to claim 1, characterized in that, Step 3 specifically involves: First, calculate the total load: （1）； Based on deformation compatibility, the total deformation of fastener 1 and connecting plate A is equal to the total deformation of fastener 2 and connecting plate B: （2）； （3）； The load on the connecting plate can then be calculated. , and the load of fasteners , ; In the formula, , This is due to the deformation of the fastener. , To accommodate the deformation of the connecting plate, The length of the connecting strip. The cross-sectional area of ​​the connecting strip is... The elastic modulus of the connecting strip plate, For the fastener's compliance constant, refer to Formula 4: （4）； In the formula, , For the thickness of the connecting plate unit; For the diameter of the fastener, , It is a constant.

3. The method for strength analysis of longitudinal connection structure of aircraft skin according to claim 1, characterized in that, The safety margin for the fasteners in step 8 is as follows: （9）； In the formula: , is the ratio of the actual tensile load of the fastener to the allowable tensile value; , is the ratio of the actual shear load of the fastener to the allowable shear value; It is the ultimate tensile strength of the fastener; It is the ultimate shear strength of the fastener.

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