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DFR fatigue calculation method for stringer through hole of transportation aircraft fuselage

A transport aircraft and long truss technology, applied in the directions of calculation, computer-aided design, geometric CAD, etc., can solve the problems that are not involved, and achieve the effect of saving time, clear theoretical basis, and simple steps

Active Publication Date: 2021-04-30
XIAN AIRCRAFT DESIGN INST OF AVIATION IND OF CHINA
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Problems solved by technology

[0004] Existing relevant patent literature does not relate to the problem described in the present invention

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  • DFR fatigue calculation method for stringer through hole of transportation aircraft fuselage
  • DFR fatigue calculation method for stringer through hole of transportation aircraft fuselage
  • DFR fatigue calculation method for stringer through hole of transportation aircraft fuselage

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Embodiment Construction

[0040] The present invention will be described in further detail below in conjunction with the accompanying drawings and examples. see Figure 1 to Figure 4 .

[0041] The present invention comprises the following steps:

[0042] Step 1. Determine the DFR value of this part according to the R angle radius of the fuselage girder passing hole;

[0043] The calculation position of the long stringer through the hole is as follows: figure 1 As shown, the fillet radius here is 6mm, which is 0.24in, check figure 2 The available DFR value is 13ksi, which is 89.6MPa.

[0044] Step 2. Determine the reference stress σ of the calculation part based on the calculation results of the finite element stress of the whole machine ref ;

[0045] The height of the girder passing holes is h=32mm, the distance between the girders is S=191mm, and the net distance between the passing holes is b=162mm; the height of the frame is h F =117mm, frame web thickness t=2.5mm, such as image 3 shown....

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Abstract

The invention belongs to the transportation aircraft structure fatigue calculation technology, and relates to DFR fatigue analysis of a transportation aircraft fuselage stringer through hole, which comprises the following steps: step 1, determining the DFR value of the part according to the R angle radius of the fuselage stringer through hole; 2, the system collects the shear flow difference between the front skin and the rear skin of the frame, and determines the reference stress sigma ref of a calculation part by combining the height of stringer passing holes, the distance between stringers, the width of angle parts, the thickness of a frame web and the height of the frame on the basis of a full-aircraft finite element stress calculation result; 3, determining the maximum stress, the minimum stress and the stress ratio in the load spectrum; 4, calculating a ground-air-ground damage ratio lambda; step 5, calculating an equivalent ground-air-ground cycle number nd; 6, calculating the allowable stress [sigma max] of the ground-air-ground cycle; and 7, calculating the fatigue margin f.

Description

technical field [0001] The invention belongs to the structural fatigue calculation technology of transport aircraft, and relates to the fatigue calculation of fuselage girder passing holes. Background technique [0002] The DFR (Detail Fatigue Rating) fatigue calculation method belongs to the nominal stress method in the fatigue analysis method, in which DFR is the inherent fatigue characteristics of the structural details themselves, and is a measure of the quality of components and the ability to withstand repeated loads. [0003] The machined frame structure of the fuselage is directly connected to the skin through the outer edge of the frame, and the long truss passes through the "rat hole" (the long stringer passes through the hole) on the machined frame, and is connected to the frame through the corner pieces. Holes are fatigue-critical parts of the fuselage structure. In the detailed design stage of the aircraft, in order to verify the fatigue strength of the long st...

Claims

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Application Information

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IPC IPC(8): G06F30/15G06F30/23G06F119/04G06F119/14
CPCG06F30/15G06F30/23G06F2119/04G06F2119/14Y02T90/00
Inventor 王亚芳史志俊闵强
Owner XIAN AIRCRAFT DESIGN INST OF AVIATION IND OF CHINA
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