CN102944211A - Method for determining area of crack arrest rib of aircraft integral wing spar - Google Patents

Method for determining area of crack arrest rib of aircraft integral wing spar Download PDF

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CN102944211A
CN102944211A CN2012104519263A CN201210451926A CN102944211A CN 102944211 A CN102944211 A CN 102944211A CN 2012104519263 A CN2012104519263 A CN 2012104519263A CN 201210451926 A CN201210451926 A CN 201210451926A CN 102944211 A CN102944211 A CN 102944211A
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rib
crack arrest
spar
wing spar
crack
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CN102944211B (en
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翟新康
秦剑波
张彦军
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Xian Aircraft Design and Research Institute of AVIC
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Abstract

The invention belongs to the field of aviation fatigue damage tolerance and particularly relates to a method for determining the area of a crack arrest rib of an aircraft integral wing spar. The method comprises the steps of: determining a bending moment M of the integral wing spar; determining a distance y from a neutral axis of an uncracked part in the integral wing spar to the crack arrest rib and an anti-bending modulus I; and calculating a crack tip stress intensity factor K when a lower edge strip of the wing spar is cracked and is expanded to the crack arrest rib and carrying out judgment. The invention discloses a method for determining structural parameters of the crack arrest rib of the integral wing spar from the field of research on the fatigue damage tolerance according to the structural form and the force transfer characteristic of the integral wing spar. The principle basis used by the method is correct; the analysis steps are clear and simple; and the calculation on section characteristics is convenient for automation iterative computation of a computer. The disclosure of the method solves the difficult problem of determining the area of the crack arrest rib of the integral wing spar.

Description

A kind of definite method of whole wing spar of airplane crack arrest rib area
Technical field
The invention belongs to aviation fatigue damage tolerance limit field, particularly relate to a kind of definite method of whole wing spar of airplane crack arrest rib area.
Background technology
Along with the extensive utilization of one-piece construction in aeroplane structure design, wing spar structural shape also begins to be changed to the design of whole wing spar (see figure 2) by combined type spar (see figure 1).This one-piece construction has reduced stress concentration portion position and potential tired source, thereby has had good anti-fatigue performance owing to reduced the use of securing member.But just owing to reduced the like this crack arrest element of " natural " of rivet hole, may have a negative impact to its damage tolerance performance.Because edge strip tension stress effect under the whole wing spar in case crackle appears in its limit, hole, just might expand to web, if crackle is too fast in the web expansion, just may have influence on Aircraft life and safety.Therefore, generally all the crack arrest rib need to be set at the web place during structural design, postpone and delay the expansion of crackle on web by the crack arrest rib, thereby guarantee that structure has lower crack growth rate and preferably crack arrest ability.And the quality of crack arrest bead structures parameter designing will directly affect the performance of its crack arrest ability.
Research about crack arrest bead structures parameter, external open source information was not found, domestic whole wing spar is carried out overtesting research, crack arrest rib position, crack arrest rib area and crack arrest rib ratio of height to thickness structural parameters were carried out research to the impact of whole wing spar damage tolerance performance, although having pointed out crack arrest rib area is to affect the most responsive structural parameters of whole wing spar damage tolerance performance, but the most responsive structural parameters (the crack arrest rib area) method of determining is not studied, failed to provide definite method of whole wing spar crack arrest rib area.
Summary of the invention
The objective of the invention is: a kind of definite method that can accurately determine the whole wing spar of airplane crack arrest rib area of crack arrest rib area is provided.
Technical scheme of the present invention is: a kind of definite method of whole wing spar of airplane crack arrest rib area may further comprise the steps:
Step 1, determine the moment M of whole wing spar: M=σ 1* s*H;
Wherein, σ 1Be edge strip axial stress under the spar, s is edge strip area under the spar, and H is the spar height;
Step 2, aspect static strength, give crack arrest rib area an initial value A 1=a 1* b 1, distance y and the composite bending modulus I of segment neutral axis to the crack arrest rib determines not ftracture in the whole wing spar; Wherein, a 1Crack arrest rib height, b 1Crack arrest rib thickness;
The axial stress σ of step 3, judgement crack arrest rib 2Whether satisfy
Figure BDA00002389749500021
Wherein, M is the moment of flexure of spar, and y does not ftracture neutral axis to the distance of crack arrest rib in the spar, and I is composite bending modulus, σ bMaterial limits intensity; If satisfy, then enter step 4, otherwise, then return step 2, give crack arrest rib area a larger value A 2=a 2* b 2, continue to judge;
Step 4, from the fracture toughness aspect, calculate the crack tip stress intensity factor K when the edge strip crack initiation extends to the crack arrest rib under the spar;
Step 5, judge whether K meets K≤K C, wherein, K CIt is the fracture toughness of material; If meet, then the crack arrest rib area of this moment meets the demands, if do not meet, then returns step 2, gives crack arrest rib area a larger value A 3=a 3* b3 ,Continue to judge, until meet the demands.
Advantage of the present invention is: the definite method that the present invention proposes a kind of whole wing spar of airplane crack arrest rib area, the present invention has provided whole wing spar crack arrest bead structures determination method for parameter according to whole wing spar version and power transmission characteristics from fatigue damage tolerance limit research field.The present invention is tightly around whole wing spar structural parameters and whole wing spar power transmission characteristics, on the basis of crack arrest rib area initialize, can satisfy simultaneously static strength, fracture toughness from crack arrest rib area, provide a kind of definite method of whole wing spar of airplane crack arrest rib area on the basis that iterates.The theoretical foundation that the present invention uses is correct, and analytical procedure is clear, simple, and section characteristic is calculated and is convenient to the computer automation iterative computation.Proposition of the present invention has solved the difficult problem of whole wing spar crack arrest rib area definition.
Description of drawings
Fig. 1 is composite spar structural profile synoptic diagram;
Fig. 2 is whole wing spar structural profile synoptic diagram;
Fig. 3 is the whole wing spar diagrammatic cross-section;
Fig. 4 is whole wing spar Crack Extension path synoptic diagram.
Wherein, edge strip under the 1-spar, edge strip on the 2-spar, 3-wing fine strain of millet web, 4-securing member, edge strip under the 5-whole wing spar, edge strip on the 6-whole wing spar, 7-whole wing spar web, 8-crack arrest rib, 9-neutral axis, W 1-lower edge strip width, t 1-lower edge strip thickness, W 2-upper edge strip width, t 2-upper edge strip thickness, t-web thickness, a 1-crack arrest rib height, b 1-crack arrest rib thickness, H are the spar height, and the non-shaded portion neutral axis is to the distance of crack arrest rib in the y-spar, and dash area is the cracking part, and other parts are the part that do not ftracture.
Embodiment
Below in conjunction with accompanying drawing the present invention is described in further details, sees also Fig. 1 to Fig. 4.
As shown in Figure 1, be composite spar structural profile synoptic diagram, composite spar is formed by connecting by securing member by edge strip, spar web on edge strip, the spar under the spar.
As shown in Figure 2, be whole wing spar structural profile synoptic diagram, whole wing spar is an integral structure component, is comprised of edge strip, whole wing spar web, crack arrest rib on edge strip, the whole wing spar under the whole wing spar.
As shown in Figure 3, for the whole wing spar diagrammatic cross-section, provided structural parameters, for.Present invention is directed at whole wing spar of airplane.
As shown in Figure 4, be spar Crack Extension path synoptic diagram, provided crackle expands to the crack arrest rib from lower edge strip crack initiation extensions path.Wherein, dash area is the cracking part, and other parts are the part that do not ftracture.
A kind of definite method of whole wing spar of airplane crack arrest rib area may further comprise the steps:
Step 1, determine the moment M of whole wing spar: M=σ 1* s*H, wherein, σ 1Be edge strip axial stress under the spar, can obtain by aircraft model that s is edge strip area under the spar, H is the spar height;
Step 2, aspect static strength, give crack arrest rib area an initial value A 1=a 1* b 1, determine that non-shaded portion structure division neutral axis is to distance y and the composite bending modulus I of crack arrest rib in the whole wing spar; Wherein, a 1Crack arrest rib height, b 1Crack arrest rib thickness; As shown in Figure 3;
The axial stress σ of step 3, judgement crack arrest rib 2Whether satisfy Wherein, M is the moment of flexure of spar, and y is that non-shaded portion structure division neutral axis is to the distance of crack arrest rib in the spar, and I is composite bending modulus, σ bMaterial limits intensity.If satisfy, then enter step 4, otherwise, then return step 2, give crack arrest rib area a larger value A 2=a 2* b 2, continue to judge;
Step 4, from the fracture toughness aspect, calculate the crack tip stress intensity factor K when the edge strip crack initiation extends to the crack arrest rib under the spar, see Fig. 4, stress intensity factor can calculate by the ANYSYS finite element analysis software;
Step 5, judge whether K meets K≤K C, wherein, K CIt is the fracture toughness of material.If meet, then the crack arrest rib area of this moment meets the demands, if do not meet, then returns step 2, gives crack arrest rib area a larger value A 3=a 3* b 3, continue to judge, until meet the demands.
Example:
The below is described in further details the present invention with a certain instantiation.
A kind of definite method of whole wing spar of airplane crack arrest rib area may further comprise the steps:
Known: H=300mm, W 1=W 2=50mm, t 1=t 2=4.5mm, a 1=15mm, b 1=4mm, t=3mm σ b=511MPa.， K C = 2277 MPa mm .
Step 1, determine the moment M of whole wing spar:
M＝σ 1*s*H＝180*50*4.5*300＝1.215×10 7Nmm；
Step 2, give crack arrest rib area an initial value A 1=15*4, non-hatched area structure division neutral axis is to the distance y=114.8mm of crack arrest rib, composite bending modulus I=4158122mm in the calculating whole wing spar 4
The axial stress of step 3, crack arrest rib
σ 2 = M * y I = 1.215 × 10 7 * 114.8 4158122 = 335.4 MPa ≤ σ b , Meet the demands;
The crack tip stress intensity factor when edge strip crack initiation extends to the crack arrest rib under step 4, the calculating spar K = 1869 MPa mm ;
Step 5, through judging, K = 1869 MPa mm ≤ K C , Meet the demands.
At this moment.Think crack arrest rib area A=15*4=60mm 2Meet the demands.
The present invention proposes a kind of definite method of whole wing spar of airplane crack arrest rib area, the present invention has provided whole wing spar crack arrest bead structures determination method for parameter according to whole wing spar version and power transmission characteristics from fatigue damage tolerance limit research field.The present invention is tightly around whole wing spar structural parameters and whole wing spar power transmission characteristics, on the basis of crack arrest rib area initialize, can satisfy simultaneously static strength, fracture toughness from crack arrest rib area, provide a kind of definite method of whole wing spar of airplane crack arrest rib area on the basis that iterates.The theoretical foundation that the present invention uses is correct, and analytical procedure is clear, simple, and section characteristic is calculated and is convenient to the computer automation iterative computation.Proposition of the present invention has solved the difficult problem of whole wing spar crack arrest rib area definition.

Claims (1)

1. definite method of a whole wing spar of airplane crack arrest rib area is characterized in that, may further comprise the steps:
Step 1, determine the moment M of whole wing spar: M=σ 1* s*H;
Wherein, σ 1Be edge strip axial stress under the spar, s is edge strip area under the spar, and H is the spar height;
Step 2, aspect static strength, give crack arrest rib area an initial value A 1=a 1* b 1, distance y and the composite bending modulus I of segment neutral axis to the crack arrest rib determines not ftracture in the whole wing spar; Wherein, a 1Crack arrest rib height, b 1Crack arrest rib thickness;
The axial stress σ of step 3, judgement crack arrest rib 2Whether satisfy
Figure FDA00002389749400011
Wherein, M is the moment of flexure of spar, and y does not ftracture neutral axis to the distance of crack arrest rib in the spar, and I is composite bending modulus, σ bMaterial limits intensity; If satisfy, then enter step 4, otherwise, then return step 2, give crack arrest rib area a larger value A 2=a 2* b 2, continue to judge;
Step 4, from the fracture toughness aspect, calculate the crack tip stress intensity factor K when the edge strip crack initiation extends to the crack arrest rib under the spar;
Step 5, judge whether K meets K≤K C, wherein, K CIt is the fracture toughness of material; If meet, then the crack arrest rib area of this moment meets the demands, if do not meet, then returns step 2, gives crack arrest rib area a larger value A 3=a 3* b 3, continue to judge, until meet the demands.
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Cited By (6)

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CN105620718A (en) * 2014-10-28 2016-06-01 中国航空工业集团公司西安飞机设计研究所 Integrated spar crack resistant rib
CN107458623A (en) * 2017-08-04 2017-12-12 中国航空工业集团公司西安飞机设计研究所 A kind of T tails aircraft vertical fin spar testpieces design method
CN107878780A (en) * 2017-11-10 2018-04-06 中国航空工业集团公司西安飞机设计研究所 A kind of method for designing wing whole wing spar parameter
CN111046610A (en) * 2019-12-26 2020-04-21 中国航空工业集团公司西安飞机设计研究所 Method for calculating dimensionless stress intensity factor of integral wing spar of airplane
CN111881608A (en) * 2020-07-31 2020-11-03 中车青岛四方机车车辆股份有限公司 Crack arrest device size obtaining method and crack arrest device
CN112784357A (en) * 2020-12-29 2021-05-11 中国航空工业集团公司西安飞机设计研究所 Parameter combination method for determining stress intensity factor of hole corner crack

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Publication number Priority date Publication date Assignee Title
CN105620718A (en) * 2014-10-28 2016-06-01 中国航空工业集团公司西安飞机设计研究所 Integrated spar crack resistant rib
CN107458623A (en) * 2017-08-04 2017-12-12 中国航空工业集团公司西安飞机设计研究所 A kind of T tails aircraft vertical fin spar testpieces design method
CN107878780A (en) * 2017-11-10 2018-04-06 中国航空工业集团公司西安飞机设计研究所 A kind of method for designing wing whole wing spar parameter
CN107878780B (en) * 2017-11-10 2020-10-09 中国航空工业集团公司西安飞机设计研究所 Method for designing parameters of integral wing spar of wing
CN111046610A (en) * 2019-12-26 2020-04-21 中国航空工业集团公司西安飞机设计研究所 Method for calculating dimensionless stress intensity factor of integral wing spar of airplane
CN111046610B (en) * 2019-12-26 2023-05-23 中国航空工业集团公司西安飞机设计研究所 Calculation method of dimensionless stress intensity factor of integral wing spar of airplane
CN111881608A (en) * 2020-07-31 2020-11-03 中车青岛四方机车车辆股份有限公司 Crack arrest device size obtaining method and crack arrest device
CN111881608B (en) * 2020-07-31 2024-02-20 中车青岛四方机车车辆股份有限公司 Method for obtaining size of crack stopper and crack stopper
CN112784357A (en) * 2020-12-29 2021-05-11 中国航空工业集团公司西安飞机设计研究所 Parameter combination method for determining stress intensity factor of hole corner crack
CN112784357B (en) * 2020-12-29 2024-02-13 中国航空工业集团公司西安飞机设计研究所 Parameter combination method for determining stress intensity factors of hole corner cracks

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