CN117034693A - Reinforced frame Liang Sunshang tolerance analysis method - Google Patents
Reinforced frame Liang Sunshang tolerance analysis method Download PDFInfo
- Publication number
- CN117034693A CN117034693A CN202310994657.3A CN202310994657A CN117034693A CN 117034693 A CN117034693 A CN 117034693A CN 202310994657 A CN202310994657 A CN 202310994657A CN 117034693 A CN117034693 A CN 117034693A
- Authority
- CN
- China
- Prior art keywords
- crack
- stress
- aircraft body
- body structure
- analysis
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
- 238000004458 analytical method Methods 0.000 title claims abstract description 74
- 230000003014 reinforcing effect Effects 0.000 claims abstract description 57
- 238000000034 method Methods 0.000 claims abstract description 47
- 230000006378 damage Effects 0.000 claims abstract description 44
- 238000001228 spectrum Methods 0.000 claims abstract description 37
- 230000008569 process Effects 0.000 claims abstract description 16
- 230000008859 change Effects 0.000 claims abstract description 8
- 238000011835 investigation Methods 0.000 claims abstract description 8
- 230000000694 effects Effects 0.000 claims abstract description 3
- 238000007689 inspection Methods 0.000 claims description 51
- 230000002787 reinforcement Effects 0.000 claims description 34
- 238000005336 cracking Methods 0.000 claims description 14
- 239000006185 dispersion Substances 0.000 claims description 14
- 239000000463 material Substances 0.000 claims description 11
- 238000004364 calculation method Methods 0.000 claims description 7
- 230000002093 peripheral effect Effects 0.000 claims description 7
- 108010063499 Sigma Factor Proteins 0.000 claims description 4
- 230000005540 biological transmission Effects 0.000 claims description 4
- 239000003351 stiffener Substances 0.000 claims description 4
- 238000013461 design Methods 0.000 abstract description 4
- 230000009471 action Effects 0.000 description 4
- 238000010586 diagram Methods 0.000 description 4
- 230000007547 defect Effects 0.000 description 3
- 238000004519 manufacturing process Methods 0.000 description 3
- 238000009826 distribution Methods 0.000 description 2
- 239000011159 matrix material Substances 0.000 description 2
- 230000035515 penetration Effects 0.000 description 2
- 208000012260 Accidental injury Diseases 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 239000012141 concentrate Substances 0.000 description 1
- 238000005260 corrosion Methods 0.000 description 1
- 230000007797 corrosion Effects 0.000 description 1
- 238000011161 development Methods 0.000 description 1
- 208000014674 injury Diseases 0.000 description 1
- 238000012423 maintenance Methods 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 230000003068 static effect Effects 0.000 description 1
- 238000012360 testing method Methods 0.000 description 1
Classifications
-
- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F30/00—Computer-aided design [CAD]
- G06F30/20—Design optimisation, verification or simulation
- G06F30/23—Design optimisation, verification or simulation using finite element methods [FEM] or finite difference methods [FDM]
-
- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
- G06F2119/02—Reliability analysis or reliability optimisation; Failure analysis, e.g. worst case scenario performance, failure mode and effects analysis [FMEA]
-
- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
- G06F2119/14—Force analysis or force optimisation, e.g. static or dynamic forces
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation
Landscapes
- Engineering & Computer Science (AREA)
- Physics & Mathematics (AREA)
- Theoretical Computer Science (AREA)
- Computer Hardware Design (AREA)
- Evolutionary Computation (AREA)
- Geometry (AREA)
- General Engineering & Computer Science (AREA)
- General Physics & Mathematics (AREA)
- Testing Of Devices, Machine Parts, Or Other Structures Thereof (AREA)
Abstract
The invention provides a method for analyzing tolerance of a reinforcing frame Liang Sunshang, and belongs to the field of structural design of large aircrafts. The method comprises the steps of determining the maximum stress position as an initial crack position after stress investigation; taking the aircraft body structure, ribs or reinforcing plates connected with the aircraft body structure and the fastening pieces as a combined structure through detail finite element analysis, and calculating the change of a crack stress intensity factor along with the crack size by considering the effect of the ribs or the reinforcing plates and the fastening pieces on cracks of the aircraft body structure so as to obtain the geometric factor of the structure; comprehensively considering the residual strength conditions of the aircraft body structure, the reinforcing frame beams and the fasteners in the crack propagation process and the structural failure condition, and correcting the stress spectrum; and performing crack propagation analysis on the aircraft body structure based on the analysis result. The invention greatly improves the analysis precision of damage tolerance and the validity of analysis results, can truly reflect the structural form and the bearing condition, and can improve the safety of the aircraft.
Description
Technical Field
The invention belongs to the field of structural design of large aircrafts, and relates to a method for analyzing tolerance of a reinforcing frame Liang Sunshang.
Background
Assessment of strength, detail design and manufacture must demonstrate that the aircraft will avoid catastrophic failure due to fatigue, corrosion, manufacturing defects or accidental injury throughout its service life, as specified by airworthiness standards. For a given structure type, it is necessary to build up its inspection threshold on the basis of crack propagation analysis and/or testing, and assume that the structure contains an initial defect of the largest size that may be caused by manufacturing or service damage.
Furthermore, the extent of damage used for the residual intensity assessment must be consistent with initial perceptibility and subsequent expansion under repeated loads at any time during the life of the vehicle, as specified by the airworthiness standards. The residual strength assessment must indicate that the remaining structure is able to withstand the corresponding situation of loading (considered as extreme static loading).
According to the overall design and functional requirements of a large-scale aircraft, the structures of the aircraft body such as a large opening of a wallboard, a reinforcing frame beam and the like are one of key points of the damage tolerance analysis of the aircraft body. The analysis must ensure that after an initial damage exists on the structure, the damage does not propagate to dangerous dimensions within a prescribed inspection cycle, and that all structures (e.g., skins, frames, beams, stringers, fasteners, etc.) still have sufficient load-bearing capacity to maintain the safety of the structure.
Damage tolerance analysis methods based on residual strength and crack growth are now well established. During development of the aircraft, corresponding methods were performed for typical structures such as flat panel longitudinal and circumferential center cracks, fastener hole corner cracks, edge corner cracks of the cap structure, and the like. However, the existing damage tolerance analysis method cannot truly reflect the combination form and bearing condition of the structure, the damage condition is not considered completely, the safety life assessment accuracy is not high, the life analysis result is too conservative, and the structure weight and the maintenance cost are greatly improved.
Disclosure of Invention
The invention aims to overcome the defects of the existing damage tolerance analysis method and provide a more accurate reinforcement frame Liang Sunshang tolerance analysis method. Based on three factors of damage tolerance, namely crack propagation analysis, residual strength analysis and a structural inspection method, the invention takes an aircraft body structure and a reinforcement, a reinforcing plate and a fastener connected with the aircraft body structure as a combined structure, determines the maximum stress position as an initial crack position after stress investigation through detail finite element analysis, and calculates the change of a crack stress intensity factor along with the crack size by considering the action of the reinforcement, the reinforcing plate and the fastener on the aircraft body structure, wherein the analysis precision is greatly superior to that of the traditional method; the method has the advantages that the residual strength conditions of the aircraft body structure, the reinforcement plates connected with the aircraft body structure and the reinforcement plates and the fasteners in the crack propagation process are comprehensively considered, the failure conditions of the aircraft body structure, the reinforcement plates connected with the aircraft body structure and the fasteners in the crack propagation process are considered, the stress spectrum is corrected, the damage tolerance analysis precision and the analysis result effectiveness are greatly improved, the structural form and the bearing condition can be reflected more truly, and the aircraft safety can be improved.
In order to achieve the above purpose, the technical scheme adopted by the invention is as follows:
a method for analyzing the tolerance of a reinforcing frame Liang Sunshang, which comprises the steps of determining the maximum stress position as an initial crack position through stress investigation; taking the aircraft body structure, ribs or reinforcing plates connected with the aircraft body structure and the fastening pieces as a combined structure through detail finite element analysis, and calculating the change of a crack stress intensity factor along with the crack size by considering the effect of the ribs or the reinforcing plates and the fastening pieces on cracks of the aircraft body structure so as to obtain the geometric factor of the structure; comprehensively considering the residual strength conditions of the aircraft body structure, the reinforcing frame beams and the fasteners in the crack propagation process, and correcting the stress spectrum by considering the failure conditions of ribs or reinforcing plates and fasteners connected with the aircraft body structure in the crack propagation process; and performing crack propagation analysis on the aircraft body structure based on the analysis result. The analysis method comprises the following specific steps:
the first step: determining a main structural part: and taking the aircraft body structure as an analysis object, and determining the material and structural form information of the aircraft body structure, wherein the information comprises geometric dimensions, skin thickness, reinforcement condition at an opening, reinforcement condition at a connecting part, model of a fastener and the like.
And a second step of: assuming initial cracking and determining cracking pattern: determining an initial crack hypothesis for the aircraft body structure, including crack location and crack size; after stress investigation, determining the maximum stress position of the aircraft body structure as an initial crack cracking position; the initial cracking mode is determined according to the aircraft body structure and the load form.
And a third step of: calculating a load spectrum and a stress spectrum: the load spectrum is a load source for analyzing the damage tolerance of the engine body structure, and the aircraft load spectrum is compiled according to the flight-continuous-flight sequence; and obtaining fatigue stress through stress analysis under fatigue load, and compiling a stress spectrum.
Fourth step: calculating stress intensity factors: establishing a finite element model of the structural details of the aircraft body, wherein the finite element model comprises the aircraft body structure, peripheral reinforcement, a reinforcing plate, fasteners and the like; carrying out finite element calculation on detail finite element models with different crack lengths, and solving stress intensity factors K of tips with different crack lengths by combining a finite element analysis result with a virtual crack closure theory VCCT method, wherein a calculation formula of the stress intensity factors K of the tips of the cracks is as follows:
wherein,
G I is the energy release rate of the I-type crack;
G II is the type II crack energy release rate;
Δa is the crack length increment;
K I is the energy release rate of the I-type crack;
K II is the type II crack energy release rate;
F x ，F y balancing the node force of the crack tip unit under a local coordinate system;
u x ,u x′ u y ,u y′ displacing the crack tip adjacent node;
e is the elastic modulus of the material;
and calculating the change of the crack stress intensity factor along with the crack size according to the formulas (1) to (4), and forming a stress intensity factor curve.
Fifth step: determining a geometric factor: combining the stress intensity factors with different crack sizes obtained from the stress intensity factor curve in the fourth step according to the formulaGeometric factors under different crack sizes are calculated, wherein beta is the geometric factor, K is the stress intensity factor, sigma is the reference working stress, namely the distal stress of the analysis structure, and c is the crack length.
Sixth step: analysis of residual intensity: calculating the stress level of the aircraft body structure under different crack sizes through a detail finite element model, and evaluating the residual strength capability of the aircraft body structure according to whether the aircraft body structure meets the residual strength load requirement and the maximum damage degree thereof, wherein the maximum damage degree is determined by analyzing the crack size under the action of the stress spectrum and the residual strength load requirement, the corresponding geometric factor and the material fracture toughness Kc; the remaining strength capabilities of the reinforcement, reinforcement panel and fastener structure were assessed in the same manner; comprehensively considering the residual strength conditions of the aircraft body structure, the peripheral reinforcement, the reinforcing plate and the fastening pieces in the crack propagation process, and correcting the stress spectrum compiled in the third step by bearing the load of the failure structure by the rest structures on the force transmission path after the partial structures fail in the crack propagation process; and re-evaluating the residual strength capacity of the aircraft body structure and the reinforcement, the reinforcing plate and the fastener structure by utilizing the corrected stress spectrum to obtain the residual strength under the condition that one or more of the aircraft body structure, the peripheral reinforcement, the reinforcing plate and the fastener fail.
Seventh step: calculating the service life of the structure: and (3) calculating the time interval between the initial crack expansion and the maximum allowable damage of the aircraft body structure according to the analysis result of the residual strength obtained in the sixth step, namely the service life of the structure.
Eighth step: determining an inspection object, an inspection method, an inspection threshold value and a repeated inspection interval: after the inspection object and the inspection method are determined, an inspection threshold value and a repeated inspection interval are calculated, wherein the inspection threshold value is obtained by dividing the structural life calculated in the seventh step by a dispersion coefficient, the repeated inspection interval is obtained by dividing the time interval from the length of the detectable crack to the critical crack length by the dispersion coefficient, the dispersion coefficient is selected according to actual conditions, and the length of the detectable crack is determined according to the inspection object and the inspection method.
Furthermore, the analysis method is suitable for damage tolerance analysis of large-opening structures of large aircraft bodies or structures of reinforced frame beams of the aircraft.
The invention has the beneficial effects that:
1. the method is used for truly simulating the real state of the body structure and crack propagation of the aircraft, has analysis precision far higher than that of the traditional analysis method, and is suitable for the analysis method of the damage tolerance of the complex body structure under various load conditions;
2. the method has the advantages that the damage tolerance analysis can be carried out on the large-opening structure of the large aircraft body, the reinforcement stress level and the reinforcement fastener load can be timely concerned, the damage tolerance analysis can be carried out on the reinforcing frame beam structure, the stress level and the fastener load of the reinforcing plate can be timely concerned, the residual strength capacity of the structure is analyzed, and the aircraft safety can be improved.
3. The damage tolerance analysis method suitable for the future commercial aircraft body structure in China is formed, and important reference basis and theoretical support are provided for large-opening structure authentication, reinforcing frame beam structure authentication, continuous seaworthiness and the like of the large-scale aircraft body.
Drawings
FIG. 1 is a schematic view of a large opening structure of the present invention.
Fig. 2 is a finite element diagram of a connection detail of a large opening structure and ribs, in the diagram, a represents an opening, B represents a crack at the opening, and C represents the ribs.
FIG. 3 is a graph of the stress intensity factor of the present invention.
Fig. 4 is a graph of the residual strength of the ribs and fastener of the present invention.
FIG. 5 is a graph of damage tolerance analysis lifetime versus large opening structure of the present invention.
FIG. 6 is a schematic diagram showing crack propagation in a reinforcement frame beam-to-reinforcement plate connection, where D represents the reinforcement frame beam, E represents the crack, and F represents the reinforcement plate.
FIG. 7 is a detailed finite element diagram of the connection of a stiffener to a stiffener frame beam structure.
FIG. 8 is a stress intensity factor graph.
Fig. 9 reinforcement plate residual strength curve.
FIG. 10 is a plot of fastener residual strength.
FIG. 11 is a graph of damage tolerance analysis life versus reinforcement frame beam structure.
Detailed Description
The technical scheme of the invention is clearly and completely described below. It is apparent that the described embodiments are only some embodiments of the present invention, not all embodiments, and that those skilled in the art may improve or adjust all other embodiments based on the embodiments of the present invention, which are within the scope of the present invention.
Example 1
Taking damage tolerance analysis of a large opening structure of an aircraft as an example, the content and the steps of a damage tolerance analysis method for the large opening structure of a large aircraft body are described with reference to the accompanying drawings, and the accompanying drawings are taken from the analysis and calculation results of the structure.
When the damage tolerance analysis of the large-opening structure of the large-scale airplane body is carried out, the basic implementation process is as follows:
the first step: determining a main structural part: for damage tolerance analysis of a large opening structure (shown in fig. 1) of a large aircraft body, the large opening structure of the large aircraft body is taken as an analysis object, and material and structural form information of the large opening structure is determined, wherein the material and structural form information comprises the geometric dimension of a large opening, the thickness of a skin, the reinforcement condition of the opening, the type of a fastener and the like.
And a second step of: assuming initial cracking and determining cracking pattern: determining an initial crack hypothesis of the large opening structure of the machine body, wherein the initial crack hypothesis comprises crack positions and crack sizes; after stress investigation, determining the maximum stress position of the opening area as an initial crack cracking position; the initial cracking mode is determined according to the structure and the load form, and the edge angle crack or the edge penetration crack is selected according to the thickness of the large-opening structure, in this embodiment, the edge angle crack, and the initial crack length is 1.27mm.
And a third step of: calculating a load spectrum and a stress spectrum: the load spectrum is a load source for analyzing the damage tolerance of the engine body structure, after high-load interception and low-load interception, a 5-by-5 spectrum load distribution matrix is established, and the aircraft load spectrum is compiled according to the flight-continuous-flight sequence; and obtaining a fatigue stress spectrum through stress analysis of a large opening structure of the machine body under fatigue load.
Fourth step: calculating stress intensity factors: the large opening structure of the large aircraft body can concentrate stress at the periphery of the opening, and the stress changes in a gradient manner along the crack propagation direction, so that a large aircraft body large opening structure detail finite element model (shown in fig. 2, wherein A is an opening, B is an opening crack, C is a rib) is built, and the large aircraft body large opening structure detail finite element model comprises an opening area skin, ribs, fasteners and the like; the skin and the reinforcement are simplified into CQUAD4 units, the position is taken as a middle plane, and the attribute is PSHELL; the fastener is simplified to a CBUSH unit. And (3) carrying out finite element calculation on detailed finite element models with different crack lengths, applying a finite element analysis result and a virtual crack closure theory VCCT method to calculate stress intensity factors K of tips with different crack lengths according to formulas (1) to (4), calculating the change of the crack stress intensity factors along with the crack sizes, and forming a stress intensity factor curve (shown in figure 3).
Fifth step: determining a geometric factor: stress intensity factors of different crack sizes obtained by combining finite element analysis are calculated according to a formulaGeometric factors under different crack sizes are calculated, wherein beta is the geometric factor, K is the stress intensity factor, sigma is the reference working stress, namely the distal stress of the analysis structure, and c is the crack length.
Sixth step: analysis of residual intensity: calculating skin stress levels under different crack sizes through a detail finite element model, and evaluating the residual strength capacity of the aircraft body structure according to whether the skin meets the residual strength load requirement and the maximum damage degree thereof, wherein the maximum damage degree is determined by analyzing the crack sizes under the action of the stress spectrum and the residual strength load requirement, the corresponding geometric factors and the material fracture toughness Kc together so as to evaluate the residual strength capacity of the skin structure; the remaining strength capabilities of the reinforcement and fastener structure were assessed in the same manner (see fig. 4). Comprehensively considering the residual strength conditions of a large opening structure, peripheral reinforcement and fasteners of the machine body in the crack propagation process, and correcting the stress spectrum compiled in the third step by bearing the load of a failure structure by the rest structures on a force transmission path after the failure of part of the structures in the crack propagation process; and re-evaluating the residual strength capability of the skin and the reinforcement and fastener structure by using the corrected stress spectrum to obtain the residual strength under the condition of one or more failure of the skin, the peripheral reinforcement and the fastener.
Seventh step: calculating the service life of the structure: according to the analysis result of the residual strength obtained in the last step, the time interval between the initial crack expansion and the maximum allowable damage of the large-opening structure of the computer body is the service life of the structure (as shown in figure 5); in this embodiment, the service life of the large opening structure of the machine body is 140816 times.
Eighth step: determining an inspection object, an inspection method, an inspection threshold value and a repeated inspection interval: after the inspection object and the inspection method are determined, an inspection threshold value and a repeated inspection interval are calculated, wherein the inspection threshold value is obtained by dividing the time interval from the initial crack propagation to the maximum allowable damage by a dispersion coefficient, and the repeated inspection interval is obtained by dividing the time interval from the detectable crack length to the critical crack length (namely, the structural life minus the corresponding flight number of the detectable crack length) by the dispersion coefficient; in this embodiment, the dispersion coefficient is 2, and the inspection threshold value is 70408 times; the length of the detectable crack is 25.4mm, the corresponding flight times are 52068 times, the dispersion coefficient is 2, and the repeated inspection interval is 44374 times.
Example 2
Taking damage tolerance analysis of a certain aircraft reinforcing frame beam structure as an example, the content and steps of a damage tolerance analysis method of the reinforcing frame beam structure are described, and the attached drawings are taken from analysis and calculation results of the structure.
The first step: determining a main structural part: for damage tolerance analysis of the reinforcing frame beam structure (in fig. 6, D is a reinforcing frame beam, E is a crack, and F is a reinforcing plate), the reinforcing frame beam structure is taken as an analysis object, and materials and structural form information of the reinforcing frame beam structure are determined, including geometric dimensions, conditions of the reinforcing plates at joints, types of fasteners and the like.
And a second step of: assuming initial cracking and determining cracking pattern: determining an initial crack hypothesis for the reinforcing frame beam structure, including crack location and crack size; and after stress investigation, determining the maximum stress position of the reinforcing frame beam as an initial crack position, wherein the initial crack mode of the reinforcing frame beam structure selects a hole corner crack or a hole edge penetration crack according to the thickness of the reinforcing frame beam structure, and the initial crack length is 1.27mm in the embodiment.
And a third step of: calculating a load spectrum and a stress spectrum: the load spectrum is a load source for analyzing the damage tolerance of the engine body structure, after high-load interception and low-load interception, a 5-by-5 spectrum load distribution matrix is established, and the aircraft load spectrum is compiled according to the flight-continuous-flight sequence; and obtaining a fatigue stress spectrum through stress analysis of the reinforced frame beam structure under fatigue load.
Fourth step: calculating stress intensity factors: stress concentration can be generated at the hole edge of the fastening piece at the joint of the reinforcing frame beam structure and the reinforcing plate, and the stress changes in a gradient manner along the crack propagation direction, so that a finite element model (shown in figure 7) of the structural details of the reinforcing frame beam is built, wherein the finite element model comprises the reinforcing frame beam, the reinforcing plate, the fastening piece and the like; the reinforced frame beams and the reinforced plates are simplified into CQUAD4 units, and the position of the reinforced frame beams and the reinforced plates is a central plane, and the property is PSHELL; the fastener is simplified into a CBUSH unit; and (3) carrying out finite element calculation on detailed finite element models with different crack lengths, applying a finite element analysis result and a virtual crack closure theory VCCT method to calculate stress intensity factors K of tips with different crack lengths according to formulas (1) to (4), calculating the change of the stress intensity factors of the cracks along with the crack sizes, and forming a stress intensity factor curve (shown in figure 8).
Fifth step: determining a geometric factor: stress intensity factors of different crack sizes obtained by combining finite element analysis are calculated according to a formulaGeometric factors under different crack sizes are calculated, wherein beta is the geometric factor, K is the stress intensity factor, sigma is the reference working stress, namely the distal stress of the analysis structure, and c is the crack length.
Sixth step: analysis of residual intensity: calculating stress levels of the reinforcing frame beams under different crack sizes through a detail finite element model, and evaluating the residual strength capability of the aircraft body structure according to whether the reinforcing frame beams meet the residual strength load requirements and the maximum damage degree of the reinforcing frame beams, wherein the maximum damage degree is determined by analyzing the crack sizes under the actions of a stress spectrum and the residual strength load requirements, corresponding geometric factors and material fracture toughness Kc; the remaining strength ability of the reinforcement panel and fastener structure was assessed in the same manner (fig. 9 and 10). Comprehensively considering the residual strength conditions of the reinforcing frame beam structure, the reinforcing plate and the fastening pieces in the crack propagation process, and correcting the stress spectrum compiled in the third step by bearing the load of the failure structure by the rest structures on the force transmission path after the partial structures fail in the crack propagation process; and re-evaluating the residual strength capability of the reinforcing frame beam, the reinforcing plate and the fastener structure by using the corrected stress spectrum to obtain the residual strength under the condition that one or more of the reinforcing frame beam, the reinforcing plate and the fastener fail.
Seventh step: calculating the service life of the structure: based on the analysis result of the residual strength obtained in the previous step, the structural life (as shown in fig. 11) which is the time interval between the initial crack propagation and the maximum allowable damage of the reinforcing frame beam structure was calculated, and in this embodiment, the structural life of the reinforcing frame beam structure was 44416 times.
Eighth step: determining an inspection object, an inspection method, an inspection threshold value and a repeated inspection interval: after the inspection object and the inspection method are determined, an inspection threshold value and a repeated inspection interval are calculated, wherein the inspection threshold value is obtained by dividing the time interval between the initial crack propagation and the maximum allowable damage (namely, the structural life minus the corresponding flight number of the detectable crack length) by a dispersion coefficient, and the repeated inspection interval is obtained by dividing the time interval between the detectable crack length and the critical crack length by the dispersion coefficient; in this embodiment, the dispersion coefficient is 2, and the inspection threshold value is 22208 times; the length of the detectable crack is 8.26mm, the corresponding flight number is 8526 times, and the dispersion coefficient is 2, and the repeated inspection interval is 17945 times.
The examples described above represent only embodiments of the invention and are not to be understood as limiting the scope of the patent of the invention, it being pointed out that several variants and modifications may be made by those skilled in the art without departing from the concept of the invention, which fall within the scope of protection of the invention.
Claims (6)
1. A method for analyzing the tolerance of a reinforcing frame Liang Sunshang, which is characterized in that the maximum stress position is determined as an initial crack position through stress investigation; taking the aircraft body structure, ribs or reinforcing plates connected with the aircraft body structure and the fastening pieces as a combined structure through detail finite element analysis, and calculating the change of a crack stress intensity factor along with the crack size by considering the effect of the ribs or the reinforcing plates and the fastening pieces on cracks of the aircraft body structure so as to obtain the geometric factor of the structure; comprehensively considering the residual strength conditions of the aircraft body structure, the reinforcing frame beams and the fasteners in the crack propagation process, and correcting the stress spectrum by considering the failure conditions of ribs or reinforcing plates and fasteners connected with the aircraft body structure in the crack propagation process; and performing crack propagation analysis on the aircraft body structure based on the analysis result.
2. The method for analyzing the tolerance of the reinforcement frame Liang Sunshang according to claim 1, comprising the specific steps of:
the first step: determining a main structural part: taking the aircraft body structure as an analysis object, and determining the material and structural form information of the aircraft body structure;
and a second step of: assuming initial cracking and determining cracking pattern: determining an initial crack hypothesis for the aircraft body structure, including crack location and crack size; after stress investigation, determining the maximum stress position of the aircraft body structure as an initial crack cracking position; the initial cracking mode is determined according to the structure and the load form of the aircraft body;
and a third step of: calculating a load spectrum and a stress spectrum: compiling an airplane load spectrum according to the flight-continuous-flight sequence; obtaining fatigue stress through stress analysis under fatigue load, and compiling a stress spectrum;
fourth step: calculating stress intensity factors: establishing a finite element model of the structural details of an aircraft body; performing finite element calculation on detail finite element models with different crack lengths, solving stress intensity factors K of tips with different crack lengths by combining a finite element analysis result with a virtual crack closure theory VCCT method, calculating the change of the crack stress intensity factors along with the crack sizes, and forming a stress intensity factor curve;
fifth step: determining a geometric factor: combining the stress intensity factors with different crack sizes obtained from the stress intensity factor curve in the fourth step according to the formulaCalculating geometric factors under different crack sizes, wherein beta is the geometric factor, K is the stress intensity factor, sigma is the reference working stress, namely the far-end stress of the analysis structure, and c is the crack length;
sixth step: analysis of residual intensity: calculating the stress level of the aircraft body structure under different crack sizes through a detail finite element model, and evaluating the residual strength capability of the aircraft body structure according to whether the aircraft body structure meets the residual strength load requirement and the maximum damage degree of the aircraft body structure; the remaining strength capabilities of the reinforcement, reinforcement panel and fastener structure were assessed in the same manner; comprehensively considering the residual strength conditions of the aircraft body structure, the peripheral reinforcement, the reinforcing plate and the fastening pieces in the crack propagation process, and correcting the stress spectrum compiled in the third step by bearing the load of the failure structure by the rest structures on the force transmission path after the partial structures fail in the crack propagation process; re-evaluating the residual strength capacities of the aircraft body structure, the reinforcement, the reinforcing plate and the fastener structure by utilizing the corrected stress spectrum to obtain the residual strength under the condition that one or more of the aircraft body structure, the peripheral reinforcement, the reinforcing plate and the fastener fail;
seventh step: calculating the service life of the structure: according to the analysis result of the residual strength obtained in the sixth step, calculating the time interval between the initial crack expansion and the maximum allowable damage of the aircraft body structure, namely the service life of the structure;
eighth step: determining an inspection object, an inspection method, an inspection threshold value and a repeated inspection interval: after the inspection object and the inspection method are determined, an inspection threshold value and a repeated inspection interval are calculated.
3. The method of claim 1, wherein in the fourth step, the crack tip stress intensity factor K is calculated as follows:
wherein,
G I is the energy release rate of the I-type crack;
G II is the type II crack energy release rate;
Δa is the crack length increment;
K I is the energy release rate of the I-type crack;
K II is the type II crack energy release rate;
F x ，F y balancing the node force of the crack tip unit under a local coordinate system;
u x ,u x′ u y ,u y′ displacing the crack tip adjacent node;
e is the elastic modulus of the material.
4. A method of analyzing the tolerance of a reinforcing frame Liang Sunshang according to claim 1, wherein in the sixth step, the maximum damage level is determined by analyzing the crack size under the stress spectrum and residual strength load requirements, the corresponding geometric factors and the material fracture toughness Kc.
5. The method according to claim 1, wherein in the eighth step, the inspection threshold value is calculated by dividing the structural life calculated in the seventh step by a dispersion coefficient, and the time interval between the length of the inspected crack and the critical crack length is divided by the dispersion coefficient at repeated inspection intervals, wherein the dispersion coefficient is selected according to the actual situation, and the length of the inspected crack is determined according to the inspection object and the inspection method.
6. A method of analyzing the tolerance of a stiffener frame Liang Sunshang according to any one of claims 1 to 5, wherein the method is applicable to the damage tolerance analysis of large open structures of large aircraft bodies or stiffener frame beam structures of aircraft.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310994657.3A CN117034693A (en) | 2023-08-09 | 2023-08-09 | Reinforced frame Liang Sunshang tolerance analysis method |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310994657.3A CN117034693A (en) | 2023-08-09 | 2023-08-09 | Reinforced frame Liang Sunshang tolerance analysis method |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN117034693A true CN117034693A (en) | 2023-11-10 |
Family
ID=88629432
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202310994657.3A Pending CN117034693A (en) | 2023-08-09 | 2023-08-09 | Reinforced frame Liang Sunshang tolerance analysis method |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN117034693A (en) |
-
2023
- 2023-08-09 CN CN202310994657.3A patent/CN117034693A/en active Pending
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| Pitt et al. | Multiple-site and widespread fatigue damage in aging aircraft | |
| Silva et al. | Multiple-site damage in riveted lap-joints: experimental simulation and finite element prediction | |
| Degenhardt et al. | Design and analysis of stiffened composite panels including post-buckling and collapse | |
| Orifici et al. | Degradation investigation in a postbuckling composite stiffened fuselage panel | |
| CN109614715B (en) | Energy field intensity method considering notch effect under action of multi-axis load and application thereof | |
| CN108984909B (en) | Mar-Lin model-based method for analyzing residual strength of composite material structure of airplane containing large damage | |
| Martin | Incorporating interlaminar fracture mechanics into design | |
| Wan et al. | Finite element modeling and fatigue life prediction of helicopter composite tail structure under multipoint coordinated loading spectrum | |
| Zamani et al. | On the influence of riveting process parameters on fatigue life of riveted lap joint | |
| CN117034693A (en) | Reinforced frame Liang Sunshang tolerance analysis method | |
| Gaitanelis et al. | Numerical FEA parametric analysis of CAI behaviour of CFRP stiffened panels | |
| Baker et al. | Repair substantiation for a bonded composite repair to F111 lower wing skin | |
| CN108491640A (en) | A kind of Multiaxial Fatigue Life Prediction model | |
| Pawar et al. | Fuzzy-logic-based health monitoring and residual-life prediction for composite helicopter rotor | |
| Goranson | Elements of structural integrity assurance | |
| CN115391947A (en) | Virtual analysis method and system for composite material structure of rail transit vehicle | |
| Newman Jr | Advances in fatigue and fracture mechanics analyses for aircraft structures | |
| Ball | The influence of residual stress on the design of aircraft primary structure | |
| Günther et al. | Composite repair for metallic aircraft structures development and qualification aspects | |
| Wetzel | Defect-tolerant structural design of wind turbine blades | |
| Viganò et al. | Investigation on the fatigue life of an Al 8090 helicopter riveted T-joint | |
| Boykett et al. | F-111C lower wing skin bonded composite repair substantiation testing | |
| Giglio et al. | Comparison of fatigue crack propagation behavior of Al 2024 and Al–Li 8090 helicopter fuselage panels | |
| Shen et al. | Composite structure design and analysis | |
| CN106769533A (en) | A kind of composite bulkhead structure four-point bending test method |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination |